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400-FORM-0002 (4/16/2014)
PACE-PWR-SPEC-0062, Revision A
Plankton, Aerosol, Cloud, ocean Ecosystem (PACE),
Code 427
PACE Solar Array Panels Specification
Goddard Space Flight Center
Greenbelt, Maryland
National Aeronautics and
Space Administration
GSFC PACE CMO
12/18/2018
Released
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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400-FORM-0002 (4/16/2014)
PACE Spacecraft Solar Array Panels Specification
Signature/Approval Page
Prepared By:
John Lyons
Reviewed By:
Beth Weinstein
Melyane Ortiz-acosta
Trevin Dear
Dave Sohl
Approved By:
Andre’ Dress
Electronic Signatures available online at: https://ipdtdms.gsfc.nasa.gov/
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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400-FORM-0002 (4/16/2014)
Preface
This document is under Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) Mission
configuration control. Changes to this document require prior approval of the PACE
Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be
submitted to the PACE Configuration Management Office (CMO), along with supportive
material justifying the proposed change. Changes to this document will be made by complete
revision.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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400-FORM-0002 (4/16/2014)
Change History Log
Revision Effective
Date
Description of Changes
(Reference the CCR & Approval Date)
Revision - 11/16/2018 Baseline Release following the approval of PACE-CCR-0459
Revision A 12/18/2018 Updated following the approval of PACE-CCR-0489
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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Table of TBDs/TBRs/TBSs
Action Item
No.
Location Summary Individual/
Organization
Actionee
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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400-FORM-0002 (4/16/2014)
Table of Contents
1.0 Introduction ........................................................................................................... 1 1.1 General Information .......................................................................................... 1 1.2 Scope ................................................................................................................ 1
2.0 Applicable Documents .......................................................................................... 2
3.0 Contract Description ............................................................................................. 4 3.1 Hardware Description ....................................................................................... 4
3.1.1 Flight Solar Array Panels ........................................................................... 4 3.1.2 Qualification Coupons ................................................................................ 4 3.1.3 Insulated Substrates .................................................................................. 4
4.0 Functional/Performance Requirements ................................................................ 5 4.1 Solar Array Panels Flight Unit Functional/Performance Requirements ............. 5
4.1.1 Test Condition Power................................................................................. 5 4.1.1.1 Flight Panels .......................................................................................... 5
4.1.1.2 Qualification Coupons ............................................................................ 5 4.1.2 Power at Highest Predicted Operating Temperature ................................. 5
4.1.3 Power at Highest Predicted Temperature .................................................. 5 4.1.4 End-of-Life Power ...................................................................................... 5 4.1.5 Solar Cell and Bypass Diode ..................................................................... 6
4.1.5.1 Solar Cell Mechanical ............................................................................. 6 4.1.5.2 Solar Cell Layout .................................................................................... 6
4.1.5.3 Solar Cell Power .................................................................................... 6 4.1.5.4 Limit to Solar Cell Shadowing ................................................................ 6
4.1.6 Solar Cell Cover ......................................................................................... 6
4.1.6.1 Cover Material and Thickness ................................................................ 6
4.1.6.2 Cover Orientation ................................................................................... 6 4.1.6.3 Antireflective/Indium Tin Oxide Coating (AR/ITO) .................................. 7
4.1.7 Blocking Diodes and Terminal Boards ....................................................... 7
4.1.8 Wire and Wire Layout ................................................................................ 7 4.1.8.1 Coarse Sun Sensor (CSS) Wiring .......................................................... 8 4.1.8.2 Hinge Damper Heater and Thermistor Wiring ........................................ 8
4.1.8.3 Monitor Solar Cell ................................................................................... 9 4.1.9 Connector Wiring and Connector Type ...................................................... 9 4.1.10 Platinum Resistor Thermometers (PRTs) ................................................ 14
4.1.11 Insulated Substrate .................................................................................. 14 4.1.11.1 Substrate Insulation Resistance ........................................................... 22 4.1.11.2 Substrate Grounding ............................................................................ 22
4.1.12 Panel Performance in Thermal Vacuum Environment ............................. 23 4.1.13 Panel Performance in Depressurization Environment ............................. 23 4.1.14 Allowable Degradation Due to Charged Particle Radiation ...................... 23 4.1.15 Allowable Degradation Due to Humidity................................................... 23
4.2 Resource Allocations ...................................................................................... 23 4.2.1 Mass Allocation ........................................................................................ 23
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4.3 Power .............................................................................................................. 23 4.3.1 Solar Array Panels Power Wire Redundancy .......................................... 23
4.4 Electrical Grounding ........................................................................................ 23 4.4.1 Primary Power DC Isolation ..................................................................... 23
4.4.2 Mechanical Contact Resistance ............................................................... 24 4.4.3 Connector DC Resistance ....................................................................... 24
4.5 Signal And Data Interfaces ............................................................................. 24 4.5.1 NA ............................................................................................................ 24 4.5.2 Passive Analog Telemetry ....................................................................... 24
4.5.2.1 Platinum Resistor Thermometer (PRT) ................................................ 24 4.5.2.2 PRT Performance................................................................................. 24
5.0 Physical Requirements ....................................................................................... 25 5.1 Interface Documentation ................................................................................. 25
5.2 Mass Properties .............................................................................................. 25 5.2.1 NA ............................................................................................................ 25
5.3 Physical Envelope ........................................................................................... 25 6.0 Environmental Requirements ............................................................................. 26
6.1 NA ................................................................................................................... 26
6.2 NA ................................................................................................................... 26 6.3 NA ................................................................................................................... 26
6.4 NA ................................................................................................................... 26 6.5 NA ................................................................................................................... 26 6.6 NA ................................................................................................................... 26
6.7 NA ................................................................................................................... 26 6.8 Pressure.......................................................................................................... 26
6.8.1 Operating Pressure Range ...................................................................... 26 6.8.2 NA ............................................................................................................ 26
6.9 NA ................................................................................................................... 26 6.10 Ground Environments ..................................................................................... 26
6.11 Thermal Requirements ................................................................................... 26 6.11.1 Flight Interface Design Temperature Limits ............................................. 26
6.12 Charged Particle Radiation Requirements ...................................................... 27 6.12.1 Definitions ................................................................................................ 27
6.12.2 Total Ionizing Dose .................................................................................. 27 6.12.2.1 Minimum TID Tolerance for EEE Parts ................................................ 27 6.12.2.2 NA ........................................................................................................ 33
6.12.3 NA ............................................................................................................ 33 6.12.4 NA ............................................................................................................ 33
6.12.4.1 NA ........................................................................................................ 33 6.12.4.2 NA ........................................................................................................ 33
6.12.4.3 NA ........................................................................................................ 33 6.12.5 Charging Environment ............................................................................. 33
6.13 Atomic Oxygen Fluence .................................................................................. 33 6.13.1 Atomic Oxygen Analysis .......................................................................... 33
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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6.13.2 Atomic Oxygen Testing ............................................................................ 33 7.0 Cleanliness ......................................................................................................... 34
7.1 Surface Contamination ................................................................................... 34 7.1.1 Surface Contamination Levels at Delivery ............................................... 34
7.1.1.1 Particulate Contamination .................................................................... 34 7.1.1.2 Molecular Contamination ...................................................................... 34
7.1.2 Surface Contamination Generation .......................................................... 34 7.1.2.1 Particulate Generation .......................................................................... 34 7.1.2.2 Molecular Generation ........................................................................... 35
7.2 Electrostatic Cleanliness ................................................................................. 35 7.2.1 Conductive Surface Ground Path ............................................................ 35 7.2.2 Conductive Surface Resistivity ................................................................ 36 7.2.3 Exposed Harness Specific Requirements ................................................ 36
7.3 Magnetic Cleanliness ...................................................................................... 36 8.0 Design & Construction Requirements ................................................................. 37
8.1 Parts, Materials & Processes (PMP) ............................................................... 37 8.1.1 EEE Parts ................................................................................................ 37 8.1.2 Materials .................................................................................................. 37
8.1.2.1 Material Conductivity ............................................................................ 37 8.1.2.2 Material Limitations for Debris Casualty Area ...................................... 37
8.2 Electrical ......................................................................................................... 38 8.2.1 Test Sensors ............................................................................................ 38 8.2.2 Interface Requirements ............................................................................ 38
8.2.2.1 Connector Selection ............................................................................. 38 8.2.2.2 Signal Segregation ............................................................................... 38
8.2.2.3 Test and Flight Signal Isolation ............................................................ 38 8.2.2.4 Test Interfaces ..................................................................................... 39
8.2.3 Mitigation of Internal Charging ................................................................. 39 8.2.3.1 NA ........................................................................................................ 39
8.2.3.2 Floating Conductors ............................................................................. 39 8.2.3.3 Dielectric Structures ............................................................................. 39
8.3 Safety .............................................................................................................. 39 8.4 NA ................................................................................................................... 40
8.5 Identification and Marking ............................................................................... 40 8.6 Workmanship .................................................................................................. 40
8.6.1 Workmanship Standards .......................................................................... 40 8.6.2 Connector ................................................................................................ 40
8.6.2.1 GSE Cable Connectors ........................................................................ 40
8.6.2.2 Prevention of Connector Mismating ..................................................... 40 8.6.2.3 NA ........................................................................................................ 41
8.6.2.4 Connector Identification ........................................................................ 41 8.6.2.5 Protection of Unused Test Connectors................................................. 41 8.6.2.6 Connector Savers................................................................................. 41
8.7 Reliability and Mission Lifetime ....................................................................... 41
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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8.7.1 Mission Life .............................................................................................. 41 8.8 Ground Handling ............................................................................................. 41
8.8.1 Ground Support Equipment (GSE) Design .............................................. 41 8.8.2 NA ............................................................................................................ 42
8.8.3 NA ............................................................................................................ 42 8.8.4 NA ............................................................................................................ 42 8.8.5 NA ............................................................................................................ 42 8.8.6 NA ............................................................................................................ 42 8.8.7 GSE Cleanliness ...................................................................................... 42
8.8.8 GSE Bakeout ........................................................................................... 42 8.8.9 Test Harness ........................................................................................... 42
9.0 Mechanical Design Requirements ...................................................................... 43 10.0 Logistics ............................................................................................................. 44
10.1 NA ................................................................................................................... 44 10.2 Ground Support Equipment ............................................................................ 44
10.3 Transportation Equipment ............................................................................... 44 11.0 Verification Requirements .................................................................................. 46
11.1 Verification Methods ....................................................................................... 46
11.1.1 Inspection ................................................................................................ 46 11.1.2 Analysis ................................................................................................... 46
11.1.3 Test .......................................................................................................... 46 11.2 Inspection Requirements ................................................................................ 46
11.2.1 Visual Inspection ...................................................................................... 46
11.2.2 Physical Measurement............................................................................. 47 11.2.3 Documentation Search ............................................................................ 47
11.3 Analysis Requirements ................................................................................... 47 11.4 Test Requirements .......................................................................................... 47
11.4.1 Definitions ................................................................................................ 47 11.4.2 NA ............................................................................................................ 48
11.4.3 Test Tolerances ....................................................................................... 48 11.4.4 Test Restrictions ...................................................................................... 48
11.4.4.1 Failure During Tests ............................................................................. 48 11.4.4.2 Modification of Hardware ...................................................................... 49
11.4.4.3 External Adjustment ............................................................................. 49 11.4.4.4 Re-Test Requirements ......................................................................... 49
11.5 Required Tests ................................................................................................ 49 11.5.1 NA ............................................................................................................ 50 11.5.2 NA ............................................................................................................ 50
11.5.3 NA ............................................................................................................ 50 11.5.4 NA ............................................................................................................ 50
11.5.5 NA ............................................................................................................ 50 11.5.6 NA ............................................................................................................ 50 11.5.7 NA ............................................................................................................ 50 11.5.8 NA ............................................................................................................ 50
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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11.5.9 NA ............................................................................................................ 50 11.5.10 Thermal Vacuum Bake-out ...................................................................... 50 11.5.11 Thermal Vacuum Test.............................................................................. 53
11.5.11.1 Thermal Vacuum Test Parameters ................................................... 53
11.5.12 NA ............................................................................................................ 55 11.5.13 NA ............................................................................................................ 55 11.5.14 NA ............................................................................................................ 55 11.5.15 NA ............................................................................................................ 55 11.5.16 Solar Cell and Bypass Diode Qualification Tests ..................................... 55
11.5.17 Solar Array Panel Qualification Tests ...................................................... 55 11.5.17.1 Solar Array Panel Life Cycle Coupon Tests ..................................... 56
11.5.18 Flight Solar Array Panel Tests ................................................................. 57 11.5.19 Test Condition Power Verification ............................................................ 57
11.5.20 Bypass Diode Functionality Verification ................................................... 57 11.5.21 Substrate Insulation Resistance Verification ............................................ 58
11.5.22 Solar Cell Mechanical Verification ........................................................... 58 11.5.23 Cover Orientation Verification .................................................................. 59 11.5.24 Cover Grounding Verification ................................................................... 59
11.5.25 Flight Connector Type Verification ........................................................... 59 11.5.26 Platinum Resistance Thermometer Type Verification .............................. 59
11.5.27 Platinum Resistance Thermometer Performance Verification .................. 59 11.5.28 Parts and Assembly Layout Verification................................................... 60 11.5.29 Mission Life Verification ........................................................................... 60
11.5.30 Shelf Life Verification ............................................................................... 60 11.5.31 Substrate Ground Verification .................................................................. 60
11.5.32 Cleanliness Verification............................................................................ 60 Appendix A Abbreviations and Acronyms .................................................................... 62
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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List of Figures Figure Page
Figure 1-1. PACE Spacecraft ......................................................................................... 1 Figure 4-1. PACE Solar Array Conceptual Diagram ....................................................... 8 Figure 4-2. Panel 3 to Panel 2 Electrical Interface ....................................................... 11
Figure 4-3. Panel 2 to Panel 1 Electrical Interface ....................................................... 12 Figure 4-4. Panel 1 to SADA Electrical Interface .......................................................... 13 Figure 4-5. Panel 1 Front and Side View ...................................................................... 15 Figure 4-6. Panel 1 Back View (KOZ = “Keep-Out Zone”) ........................................... 16 Figure 4-7. Panel 2 Front and Side View ...................................................................... 17
Figure 4-8. Panel 2 Back View (KOZ = “Keep-Out Zone”) ........................................... 18 Figure 4-9 Panel 3 Front and Side View........................................................................ 19
Figure 4-10. Panel 3 Back View (KOZ = “Keep-Out Zone”) ......................................... 20 Figure 4-11. Qualification Coupon Front and Side View............................................... 21
Figure 4-12. Qualification Coupon Back View (KOZ = “Keep-Out Zone”) .................... 22 Figure 6-1. Total Ionizing Dose-Depth Curve (includes x2 margin) .............................. 28
List of Tables Table Page
Table 2-1. Applicable Documents .................................................................................. 2 Table 6-1. Solar Array Panel Temperature Environment .............................................. 27 Table 6-2. Dose (including x2 margin) as a Function of Shielding ............................... 29
Table 6-3. Trapped Electron Spectrum ........................................................................ 30
Table 6-4. Trapped Proton Spectrum ........................................................................... 31 Table 6-5. Solar Proton Spectrum ................................................................................ 32 Table 8-1. Limited materials for debris casualty area ................................................... 37
Table 11-1. Test Tolerances ........................................................................................ 48
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
Effective Date: December 18, 2018
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400-FORM-0002 (4/16/2014)
1.0 INTRODUCTION
1.1 GENERAL INFORMATION
The Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) mission is a strategic climate continuity
mission that will extend the high quality ocean ecological, ocean biogeochemical, cloud, and
aerosol particle data records begun by NASA in the 1990s. The mission will be capable of
collecting radiometric and polarimetric measurements of the ocean and atmosphere, from which
these biological, biogeochemical, and physical properties will be determined. PACE data
products will not only add to existing critical climate and Earth system records, but also answer
new and emerging advanced science questions related to Earth’s changing climate.
An artist’s conception of the PACE spacecraft is presented in Figure 1-1.
Figure 1-1. PACE Spacecraft
1.2 SCOPE
This specification describes the electrical, mechanical, environmental, and verification testing
requirements for space-qualified Solar Array Panels that will provide electric power for the
NASA Goddard Space Flight Center (GSFC) PACE Mission.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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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)
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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
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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.
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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.
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4.1.5 Solar Cell and Bypass Diode
4.1.5.1 Solar Cell Mechanical
No cell on a panel shall have a crack, visible at 7x or less magnification.
4.1.5.2 Solar Cell Layout
a. The Contractor shall propose the solar cell layout. The solar cell layout of the three solar
array panels shall be divided into 24 parallel segments, each having an equal number of
solar cell strings in parallel, except for one segment, which may have a smaller number of
strings if the total power generated by the sum of all the parallel segments is sufficient to
meet the test condition power requirement in 4.1.1. At BOL, 1.05 AM0, 115°C, the short-
circuit current of any segment shall not exceed 3.8 amperes.
Flight and qualification panel dimensions including stay-out zones are included in
Figures 4-5 through 4-12.
b. To the extent practicable, the Contractor shall not divide individual segments between
panels.
4.1.5.3 Solar Cell Power
At 28°C and Air Mass Zero, the PACE solar cells shall produce sufficient power to meet the test
condition power specified in Section 4.1.1.
4.1.5.4 Limit to Solar Cell Shadowing
No cell shall experience more than 3 percent degradation over the life of the mission in
maximum power output as a result of the cell being repeatedly shadowed.
4.1.6 Solar Cell Cover
4.1.6.1 Cover Material and Thickness
Each coverglass shall be 100 µm (nominal) thick cerium dioxide doped glass (CMG or
equivalent) and shall cover 100 percent of the active area of each solar cell.
4.1.6.2 Cover Orientation
The contractor shall orient the cover using either an etch symbol or a stain.
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4.1.6.3 Antireflective/Indium Tin Oxide Coating (AR/ITO)
The contractor shall use covers that are coated with AR/ITO having a nominal resistivity of 109
ohms per square with the coating on each cover grounded to the solar cell front contact pad.
4.1.7 Blocking Diodes and Terminal Boards
a. The Contractor shall design the terminal (sometimes called diode) boards.
b. The Contractor shall insulate the terminal boards from and bond them to the rear face
sheet with an adhesive or hardware proposed by the Contractor and approved by the
GSFC.
c. Each terminal board shall contain parallel redundant blocking diodes which connect in
series with the solar cell strings.
d. The contractor shall conformally coat the terminal boards prior to delivery to GSFC.
4.1.8 Wire and Wire Layout
Figure 4-1 is a conceptual diagram depicting components that are to be wired to the panel
connectors.
a. For the Panel 3 to Panel 2 power and signal harnesses, the contractor shall provide 90-cm
pigtails from the outboard edge of Panel 2. These will be formed into shape at GSFC and
mated to the Panel 3 inboard connectors in accordance with the pin assignments to be
defined by the contractor.
b. Similarly, for the Panel 2 to Panel 1 power and signal harnesses, the contractor shall
provide 90-cm pigtails from the outboard edge of Panel 1. These will be formed into
shape at GSFC and mated to the Panel 2 inboard connectors in accordance with the pin
assignments to be defined by the contractor.
c. The contractor shall protect wire wherever abrasion may be a problem.
d. The contractor shall use stress relief between wire tie points to avoid strains, particularly
on the solar cell string terminations.
e. The contractor shall address how it will stake wire.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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Figure 4-1. PACE Solar Array Conceptual Diagram
4.1.8.1 Coarse Sun Sensor (CSS) Wiring
a. The contractor shall route wiring for the Coarse Sun Sensors (CSSs) to the panel
connectors on the backs of the panels using GFE twisted, shielded wire.
b. The Contractor shall ground the shield to the connector shell.
c. NASA/GSFC will attach the CSSs after the panels have been delivered. To provide for
this, the Contractor shall terminate the CSS wiring in a 30 cm pigtail at the CSS locations
on the outboard panel.
4.1.8.2 Hinge Damper Heater and Thermistor Wiring
a. The contractor shall route wiring for the hinge damper heaters to the panel connectors on
the backs of the panels using twisted wire.
Panel 3 Panel 2 Panel 1
Segment 1
Segment 8
SubstrateGround
SubstrateGround
SubstrateGround
SubstrateGround
SubstrateGround
SubstrateGround
Segment 9
Segment 16
Segment 17
Segment 24
Pos. Plus Rtn. Wire Pair
Twisted, Shielded Pair
Segment Incl. Terminal Boards, Blocking Diodes, Test Connectors
Ground Wire
1. All pos. wires are run adjacent to their returns to the extent practicable.2. PRT, CSS, and Hinge Pot wires twisted and shielded. Shield grounded to connector shell.
Connector
3. This conceptual diagram does not show physical locations of components.4. Most segments not depicted. Their wires would run in a manner similar to those shown.
Coarse Sun Sensor (CSS)
PRT
PRTPRT
PRT
Hinge Damper w/Heater
PRT
PRT
SAD
A In
terface
Monitor Solar Cell
Twisted, Shielded Trio
Thermistor Thermistor
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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.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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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.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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400-FORM-0002 (4/16/2014)
Figure 4-2. Panel 3 to Panel 2 Electrical Interface
Substrate Ground
Segment 17 Pos.
Segment 18 Pos.
Segment 19 Pos.
Segment 20 Pos.
Segment 21 Pos.
Segment 22 Pos.
Segment 23 Pos.
Segment 24 Pos.
Segment 17 Rtn.
Segment 18 Rtn.
Segment 19 Rtn.
Segment 20 Rtn.
Segment 21 Rtn.
Segment 22 Rtn.
Segment 23 Rtn.
Segment 24 Rtn.
Substrate Ground
Segment 17 Pos.
Segment 18 Pos.
Segment 19 Pos.
Segment 20 Pos.
Segment 21 Pos.
Segment 22 Pos.
Segment 23 Pos.
Segment 24 Pos.
Segment 17 Rtn.
Segment 18 Rtn.
Segment 19 Rtn.
Segment 20 Rtn.
Segment 21 Rtn.
Segment 22 Rtn.
Segment 23 Rtn.
Segment 24 Rtn.
Panel 3 Panel 2
Notes1. All wires except substrate ground and shields go through to the Panel 1-SADA interface. 2. Shields are grounded to connector shells. 3. Connector shells are grounded to the substrate honeycomb.4. Substrate ground wires are imbedded in honeycomb cells on all three panels.5. The contractor shall define the connector pinouts per the requirements of this specification.6. The contractor shall furnish the outboard side of Panel 2 with a pigtail per this sepcification. The pigtail will be made into an inter-panel harness at GSFC and inserted into connectors on the inboard side of Panel 3 per the pin assignments defined by the contractor.
CSS 1 Pos.
CSS 2 Pos.
CSS 3 Pos.
PRT 5 Pos.
PRT 6 Pos.
CSS 1 Rtn.
CSS 2 Rtn.
CSS 3 Rtn.
PRT 5 Rtn.
PRT 6 Rtn.CSS 1 Shield
CSS 2 Shield
CSS 3 Shield
PRT 5 Shield
PRT 6 Shield
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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400-FORM-0002 (4/16/2014)
Figure 4-3. Panel 2 to Panel 1 Electrical Interface
Substrate Ground
Segment 9 Pos.
Segment 10 Pos.
Segment 11 Pos.
Segment 12 Pos.
Segment 13 Pos.
Segment 14 Pos.
Segment 15 Pos.
Segment 16 Pos.
Segment 9 Rtn.
Segment 10 Rtn.
Segment 11 Rtn.
Segment 12 Rtn.
Segment 13 Rtn.
Segment 14 Rtn.
Segment 15 Rtn.
Segment 16 Rtn.
Panel 2 Panel 1
Notes1. All wires except substrate ground and shields go through to the Panel 1-SADA interface. 2. Shields are grounded to connector shells. 3. Connector shells are grounded to the substrate honeycomb.4. Substrate ground wires are imbedded in honeycomb cells on all three panels.5. The contractor shall define the connector pinouts per the requirements of this specification.6. The contractor shall furnish the outboard side of Panel 1 with a pigtail per this sepcification. The pigtail will be made into an inter-panel harness at GSFC and inserted into connectors on the inboard side of Panel 2 per the pin assignments defined by the contractor.
PRT 3 Pos.
PRT 4 Pos.
PRT 3 Rtn.
PRT 4 Rtn.
PRT 3 Shield
PRT 4 Shield
Substrate Ground
Segment 9 Pos.
Segment 10 Pos.
Segment 11 Pos.
Segment 12 Pos.
Segment 13 Pos.
Segment 14 Pos.
Segment 15 Pos.
Segment 16 Pos.
Segment 9 Rtn.
Segment 10 Rtn.
Segment 11 Rtn.
Segment 12 Rtn.
Segment 13 Rtn.
Segment 14 Rtn.
Segment 15 Rtn.
Segment 16 Rtn.
Hinge Damper Heater 2 Pos.
Hinge Damper Heater 2 Rtn.
Hinge Damper 2 Thermistor Shield
Hinge Damper 2 Thermistor Rtn.
Hinge Damper 2 Thermistor Pos.
Power and Signal Wires from Panel 3
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Figure 4-4. Panel 1 to SADA Electrical Interface
Substrate Ground
Segment 1 Pos.
Segment 2 Pos.
Segment 3 Pos.
Segment 4 Pos.
Segment 5 Pos.
Segment 6 Pos.
Segment 7 Pos.
Segment 8 Pos.
Segment 1 Rtn.
Segment 2 Rtn.
Segment 3 Rtn.
Segment 4 Rtn.
Segment 5 Rtn.
Segment 6 Rtn.
Segment 7 Rtn.
Segment 8 Rtn.
Panel 1 SADA
Notes1. All wires except shields go through to the Panel 1-SADA interface. 2. Shields are grounded to connector shells. 3. Connector shells are grounded to the substrate honeycomb.4. Substrate ground wires are imbedded in honeycomb cells on all three panels.5. The contractor shall define the connector pinouts per the requirements of this specification.
PRT 1 Pos.
PRT 2 Pos.
PRT 1 Rtn.PRT 2 Rtn.
PRT 1 Shield
PRT 2 Shield
Hinge Damper Heater 1 Pos.
Hinge Damper Heater 1 Rtn.
Hinge Damper 1 Thermistor Shield
Hinge Damper 1 Thermistor Rtn.
Hinge Damper 1 Thermistor Pos.
Power and Signal Wires from Panels 2 and 3
Substrate Ground
Segment 1 Pos.
Segment 2 Pos.
Segment 3 Pos.
Segment 4 Pos.
Segment 5 Pos.
Segment 6 Pos.
Segment 7 Pos.
Segment 8 Pos.
Segment 1 Rtn.
Segment 2 Rtn.
Segment 3 Rtn.
Segment 4 Rtn.
Segment 5 Rtn.
Segment 6 Rtn.
Segment 7 Rtn.
Segment 8 Rtn.
Monitor Solar Cell Pos.
Monitor Solar Cell Rtn.
Monitor Solar Cell Shield
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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.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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Figure 4-5. Panel 1 Front and Side View
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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Figure 4-6. Panel 1 Back View (KOZ = “Keep-Out Zone”)
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Figure 4-7. Panel 2 Front and Side View
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
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Figure 4-8. Panel 2 Back View (KOZ = “Keep-Out Zone”)
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Figure 4-9 Panel 3 Front and Side View
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Figure 4-10. Panel 3 Back View (KOZ = “Keep-Out Zone”)
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Figure 4-11. Qualification Coupon Front and Side View
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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.
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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.
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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.
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400-FORM-0002 (4/16/2014)
5.0 PHYSICAL REQUIREMENTS
5.1 INTERFACE DOCUMENTATION
The contractor shall use metric units when interfacing with NASA GSFC including any
drawings, documents, models, except for the following cases:
Heritage Hardware: Hardware that has been previously qualified, or of similar design
heritage, may be specified in English units where use of metric equivalents would
lead to additional cost to the program.
Fasteners: Although bolt patterns will be defined using metric dimensioning, use of
English fasteners (with hole dimensioning and tolerancing) is permitted.
The dimensions of the GFE insulated substrates including stay-out zones are defined in Figures
4-1 through 4-6.
The electrical interface is defined in this Specification.
5.2 MASS PROPERTIES
5.2.1 NA
5.3 PHYSICAL ENVELOPE
No Solar Array Panel shall exceed the thermal and mechanical volume envelope described in
Figures 4-1 through 4-6.
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400-FORM-0002 (4/16/2014)
6.0 ENVIRONMENTAL REQUIREMENTS
Environmental design requirements for the spacecraft components are specified in this section.
The SA Panels shall meet its performance requirements in Section 4.0 during and after exposure
to the environments specified in this section.
6.1 NA
6.2 NA
6.3 NA
6.4 NA
6.5 NA
6.6 NA
6.7 NA
6.8 PRESSURE
6.8.1 Operating Pressure Range
The Solar Array Panels shall be designed to meet all performance requirements while operating
over a pressure range of 1.08 x 105 Pa (813 Torr) to 1.3 x 10-12 Pa (1 x 10-14 Torr).
6.8.2 NA
6.9 NA
6.10 GROUND ENVIRONMENTS
The Solar Array Panels shall meet all of their performance requirements during exposure to air
temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.
6.11 THERMAL REQUIREMENTS
6.11.1 Flight Interface Design Temperature Limits
The solar array panels shall meet the requirements of this document after exposure to the
temperature extremes and number of eclipse cycles in Error! Reference source not found..
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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..
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Figure 6-1. Total Ionizing Dose-Depth Curve (includes x2 margin)
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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
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Table 6-3. Trapped Electron Spectrum
Energy
(>MeV)
Integral
Fluence
(e/cm2)
0.001 4.416E+14
0.0013 4.178E+14
0.0017 3.935E+14
0.0021 3.757E+14
0.0028 3.520E+14
0.0036 3.314E+14
0.0046 3.110E+14
0.0059 2.893E+14
0.0077 2.644E+14
0.01 2.391E+14
0.013 2.130E+14
0.016 1.936E+14
0.021 1.733E+14
0.027 1.603E+14
0.035 1.516E+14
0.04 1.483E+14
0.07 9.981E+13
0.1 7.332E+13
0.25 2.767E+13
0.5 8.377E+12
0.75 3.175E+12
1 1.461E+12
1.5 3.884E+11
2 1.108E+11
2.5 3.907E+10
3 1.711E+10
3.5 7.712E+09
4 3.375E+09
4.5 1.307E+09
5 3.786E+08
5.5 7.789E+07
6 1.022E+07
6.5 0.000E+00
7 0.000E+00
7.5 0.000E+00
8 0.000E+00
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Table 6-4. Trapped Proton Spectrum
Energy
(>MeV)
Integral
Fluence
(p/cm2)
Energy
(>MeV)
Integral
Fluence
(p/cm2)
0.00115 8.302E+13 6 1.636E+11
0.0021 7.756E+13 8 1.199E+11
0.0037 7.037E+13 10 9.614E+10
0.0065 6.194E+13 15 6.790E+10
0.01155 5.266E+13 20 5.530E+10
0.0204 4.268E+13 30 4.442E+10
0.036 3.107E+13 50 3.347E+10
0.06375 1.705E+13 60 3.016E+10
0.085 9.268E+12 80 2.533E+10
0.1 4.917E+12 100 2.153E+10
0.2 3.626E+12 150 1.447E+10
0.4 1.992E+12 200 9.695E+09
0.6 1.232E+12 300 4.056E+09
0.8 9.084E+11 400 1.637E+09
1 7.678E+11 700 1.912E+08
2 4.749E+11 1200 1.101E+07
4 2.523E+11 2000 2.000E+00
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Table 6-5. Solar Proton Spectrum
Energy
(>MeV)
Integral
Fluence
(p/cm2)
Energy
(>MeV)
Integral
Fluence
(p/cm2)
Energy
(>MeV)
Integral
Fluence
(p/cm2)
0.1 5.045E+11 2 5.693E+10 55 9.945E+08
0.11 4.711E+11 2.2 5.314E+10 63 7.684E+08
0.12 4.423E+11 2.5 4.842E+10 71 6.068E+08
0.14 3.951E+11 2.8 4.453E+10 80 4.752E+08
0.16 3.582E+11 3.2 4.004E+10 90 3.701E+08
0.18 3.287E+11 3.5 3.708E+10 100 2.948E+08
0.2 3.045E+11 4 3.295E+10 110 2.406E+08
0.22 2.841E+11 4.5 2.970E+10 120 2.002E+08
0.25 2.588E+11 5 2.700E+10 140 1.421E+08
0.28 2.383E+11 5.5 2.464E+10 160 1.043E+08
0.32 2.162E+11 6.3 2.149E+10 180 7.911E+07
0.35 2.026E+11 7.1 1.900E+10 200 6.159E+07
0.4 1.838E+11 8 1.670E+10 220 4.889E+07
0.45 1.686E+11 9 1.465E+10 250 3.540E+07
0.5 1.562E+11 10 1.300E+10 280 2.629E+07
0.55 1.458E+11 11 1.160E+10 320 1.828E+07
0.63 1.320E+11 12 1.041E+10 45 1.427E+09
0.71 1.210E+11 14 8.554E+09 50 1.183E+09
0.8 1.109E+11 16 7.171E+09 55 9.945E+08
0.9 1.018E+11 18 6.103E+09 320 1.828E+07
1 9.430E+10 25 3.778E+09 350 1.435E+07
1.1 8.801E+10 28 3.165E+09 400 9.895E+06
1.2 8.264E+10 32 2.549E+09 450 7.068E+06
1.4 7.383E+10 35 2.201E+09 500 5.295E+06
1.6 6.696E+10 45 1.427E+09
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1.8 6.145E+10 50 1.183E+09
6.12.2.2 NA
6.12.3 NA
6.12.4 NA
6.12.4.1 NA
6.12.4.2 NA
6.12.4.3 NA
6.12.5 Charging Environment
The Solar Array Panels shall be designed to withstand the degradation of surface materials and
associated surface charging effects due to the radiation environment for the PACE mission orbit.
6.13 ATOMIC OXYGEN FLUENCE
When a component has surfaces that will be exposed to the external space environment, all external
materials shall survive an atomic oxygen fluence (Observatory orbit velocity direction) of 2.0E+20
atoms/cm2 without loss of structural integrity or loss of critical performance criteria.
6.13.1 Atomic Oxygen Analysis
An analysis shall be performed for all external materials and finishes to verify the compatibility with
the AO environment. The analysis should show that any AO degradation does not pose a
contamination hazard for other components on the spacecraft (source of particles, molecular films,
debris, etc.).
6.13.2 Atomic Oxygen Testing
If no data exists for the proposed material and finishes, a test shall be performed, exposing a
representative sample to AO and verifying no loss of structural integrity or loss of critical
performance criteria.
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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.).
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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.
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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.
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8.0 DESIGN & CONSTRUCTION REQUIREMENTS
8.1 PARTS, MATERIALS & PROCESSES (PMP)
8.1.1 EEE Parts
See the requirements in the SOW, PACE-PWR-SOW-0025.
8.1.2 Materials
See the requirements in the SOW, PACE-PWR-SOW-0025.
8.1.2.1 Material Conductivity
All parts should be passivated and mounting surfaces on Solar Array Panels shall be conductive
as defined in Section 4.4.
8.1.2.2 Material Limitations for Debris Casualty Area
a. Aluminum parts shall be finished with iridite per MIL-DTL-5541, Class 3.
b. Titanium surfaces shall be finished per AMS 2488.
a. The contractor shall report the mass, dimensions, and use of materials shown in Error!
Reference source not found.. Materials should be assessed using the total quantity
within that assembly. Isolated pieces of material may weigh less than 50 grams; however,
if attached to other parts of the same material it will stay together until demise or until
reaching ground.
Table 8-1. Limited materials for debris casualty area
Material Quantity
Titanium Mass > 50g
Steel Mass > 50g
Inconel Mass > 50g
Invar Mass > 50g
Materials with melting
points above 1200° C
(1473 K)
Mass > 50g
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8.2 ELECTRICAL
8.2.1 Test Sensors
With the exception of thermocouples used in thermal ambient and thermal vacuum cycling, and
in high temperature electrical measurements, test sensors shall be designed for flight. Unless
specified to be removed before flight, test sensors will not be removed prior to flight.
8.2.2 Interface Requirements
8.2.2.1 Connector Selection
8.2.2.1.1 Connector Specifications
a. Selected connector types shall meet the Connector and Contact Requirements defined in
Section C2 of EEE-INST-002.
b. Environmental seals shall not be used in connectors especially if made from silicone
without the explicit approval by the NASA/GSFC COR.
8.2.2.1.2 NA
8.2.2.1.3 Contact Derating
The current carrying capacity of the contacts shall be derated for continuous operation at the
required current levels in a vacuum, as defined in Section C2 of EEE-INST-002.
8.2.2.1.4 Redundant Contact Derating
When redundant contacts are used for a single power source, each contact shall meet the
required derating criteria.
8.2.2.2 Signal Segregation
a. Wherever possible, different classes of signals (power, digital, analog, etc.) shall be
separated by using separate connectors and separate harness bundles.
b. If separate connectors are not feasible, classes of signals within a common connector
shall be isolated from one another. Connector pin assignments should be such that
sensitive circuits are separated from potential interference sources.
8.2.2.3 Test and Flight Signal Isolation
Test signals and flight signals shall not be located in the same connector.
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8.2.2.4 Test Interfaces
Component test signals that require access during observatory-level testing will be handled as
follows:
8.2.2.4.1 Facility-Induced Noise
All test signals should be protected or isolated from facility-induced noise.
8.2.2.4.2 Facility-Induced ESD GSE Malfunction
All test signals shall be protected or isolated from facility-induced ESD GSE malfunction.
8.2.2.4.3 Facility-Induced GSE Malfunction
All test signals shall be protected or isolated from facility-induced GSE malfunction.
8.2.3 Mitigation of Internal Charging
Internal charging refers to the physical effect where high energy electrons deposit charge in a
dielectric, if the charging rate is higher than the leakage rate eventually a point is reached where
the dielectric discharges to the nearby structure.
8.2.3.1 NA
8.2.3.2 Floating Conductors
Floating conductors, if present, shall have a bleed path of less than 10 MΩ to the component
structure. This requirement is not applicable to small floating conductors (1 inch2 or less) or short
(1 inch or less) unterminated traces or wires that are inside of the components.
8.2.3.3 Dielectric Structures
8.2.3.3.1 Bulk Resistivity
Dielectric structures shall have a bulk resistivity less than 1012 ohm-cm.
8.2.3.3.2 Charge Bleed-Off
All dielectric structures shall have a charge bleed path to the spacecraft interface, designed to
route the discharge into the spacecraft structure in a controlled fashion.
8.3 SAFETY
All flight components shall meet the applicable sections of NASA-STD-8719.24 ELV Safety
Requirements.
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a. All Ground Support Equipment being used at NASA/GSFC shall meet NASA
requirements specified in Section 8.8.
b. All Ground Support Equipment being used at the Launch site shall meet the applicable
sections of NASA-STD-8719.24 ELV Safety Requirements.
8.4 NA
8.5 IDENTIFICATION AND MARKING
Each unit shall be permanently marked with the part number and a unique sequential serial
number in the area designated on the interface control drawing in a manner to be approved by the
NASA/GSFC COR.
All markings shall use alcohol proof ink, engraving, or laser etching.
8.6 WORKMANSHIP
8.6.1 Workmanship Standards
See the workmanship standards and processes outlined in the SOW.
8.6.2 Connector
8.6.2.1 GSE Cable Connectors
GSE cable connectors that mate with flight test connectors shall be flight-approved connectors.
8.6.2.2 Prevention of Connector Mismating
Connector mismating prevention requirements are identified in this section.
8.6.2.2.1 Connector Uniqueness
Physically adjacent connectors shall be of different sizes or of different sexes or uniquely keyed
to facilitate proper mating.
8.6.2.2.2 NA
8.6.2.2.3 NA
8.6.2.2.4 Accessibility
The Solar Array Panels Spacecraft interface connectors shall be spaced far enough apart to allow
the mate and demate operations to be performed without a special tool.
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8.6.2.2.5 Connector Gender
The connector half that sources power to another Solar Array Panel shall be female (socketed) to
protect against inadvertent grounding prior to mating.
8.6.2.3 NA
8.6.2.4 Connector Identification
Each connector shall be labeled and clearly visible to facilitate proper mating.
8.6.2.5 Protection of Unused Test Connectors
Test connectors shall be capped with flight-approved RF and static control covers when not in
use.
8.6.2.6 Connector Savers
Connector savers shall be used during integration and test to minimize wear on connector
contacts and must meet the same requirements as the flight connectors.
8.7 RELIABILITY AND MISSION LIFETIME
8.7.1 Mission Life
The Solar Array Panels shall meet all performance specifications through two (2) years of
ground testing and three (3) years of operation in space, following a two (2) month
commissioning period.
8.8 GROUND HANDLING
8.8.1 Ground Support Equipment (GSE) Design
All electrical EGSE or support equipment shall be in compliance with the National Electric Code
(NFPA 70) or equivalent standard.
Do not lift items heavier than you can comfortably lift. The National Institute for Occupational
Safety and Health (NIOSH) lifting guidelines state that a person could lift 37.65 lb (17.1 kg)
under ideal conditions in front of the body not involving trunk twisting. Ideal conditions include:
smooth lifting (no jerking), the hands spread 30 inches (76 cm) or less, lifting posture
unrestricted, and object held close to the body. Additionally, there must be good couplings
(handles, shoes, floor surfaces) and a favorable environment. If all of these conditions are not
met, the maximum weight must be decreased (or alternative lifting methods should be
considered)
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8.8.2 NA
8.8.3 NA
8.8.4 NA
8.8.5 NA
8.8.6 NA
8.8.7 GSE Cleanliness
All Ground and Test support equipment shall be compatible with the flight component and the
environment where the flight component or test component will reside (cleanroom, thermal
vacuum chamber, vibration cell, etc.)
8.8.8 GSE Bakeout
Thermal Vacuum GSE shall be baked out and the outgassing rate certified prior to the test.
8.8.9 Test Harness
Test harnesses that will be used in vacuum during ground operations shall be vacuum
compatible.
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9.0 MECHANICAL DESIGN REQUIREMENTS
For additional guidance on the design and analysis of threaded fastening systems in NASA
spaceflight hardware, consult NASA-STD-5020, Requirements for Threaded Fastening Systems
In Spaceflight Hardware.
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10.0 LOGISTICS
10.1 NA
10.2 GROUND SUPPORT EQUIPMENT
a. Any GSE that will be used in a Thermal Vacuum test (harnesses, fixtures, stimulators,
etc.) shall be fabricated of vacuum compatible materials that meet the requirements of
Section 7.1.
b. Any GSE that will be used in a Thermal Vacuum test shall be capable of being baked out
at 115 degrees C. Special care will be taken to preserve the surface cleanliness of
thermal vacuum GSE items (especially harnesses) during integration and test activities in
non-cleanroom areas. Minimize contact with contaminating surfaces. Bagging is
recommended whenever possible.
c. GSE that will be used in contractor or NASA/GSFC cleanrooms shall be cleanroom
compatible.
d. Materials shall not outgas or generate particles that violate the requirements of Section
7.1.
e. Aluminum, for example, shall have protective finishes to prevent oxidation (anodize,
chromate conversion (irridite), or cleanroom compatible paint).
f. Any GSE that will be in contact or close proximity (up to approx. 6 feet) of flight
hardware shall meet the same surface cleanliness requirements as the flight hardware
unless the flight hardware can be thoroughly cleaned to remove any transferred
contaminants.
g. Materials that contact the flight hardware shall meet the vacuum compatibility
requirements in Section 7.1.
10.3 TRANSPORTATION EQUIPMENT
a. Transportation equipment shall assure that the solar array panels meet the requirements
of this specification after shipment.
b. Materials that are exposed to the same air that surrounds the flight panels, shall not
outgas hydrocarbons or other contaminants that can contaminate the panels. Bagging
materials for such large assemblies seldom provide 100 percent vapor barriers.
c. Unless the vendor can demonstrate 100 percent vapor barrier construction, the vendor
shall assure that all materials that are interior to a shipping container will not outgassing
during transport.
d. The containers shall maintain the cleanliness of the panels during long term storage – up
to 2 years.
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e. Materials of construction as well as equipment in the transportation container shall not
degrade the surface or outgassing cleanliness of the solar panels.
Materials of concern include elastomeric seals, paints, surface finishes, substrates, elastomeric
dampeners that can outgas, lubricants, wire spring dampeners with lubricants to prevent
corrosion), monitoring equipment, surface finishes, etc. Where possible materials should meet
flight outgassing requirements. Containers should be compatible with a cleanroom and capable
of being cleaned with Isopropyl alcohol. Other materials may be used but tests should be run on
them to determine their outgassing potential in air over at the high end of the shipping
temperature range. Silicones should be avoided. Wood should not be used with prior approval of
the NASA/GSFC COR and review of contamination risk mitigation strategies.
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11.0 VERIFICATION REQUIREMENTS
The contractor shall conduct a verification program that demonstrates the component design is
qualified. Per the SOW (PACE-PWR-SOW-0025), the contractor will provide a verification
matrix defining the method of verification for each specific requirement of this document.
11.1 VERIFICATION METHODS
Verification methods include inspection, analysis, as well as environmental, functional, and
performance testing, or a combination of these techniques.
11.1.1 Inspection
Verification by inspection includes (but is not limited to) visual inspection, simple physical
manipulation, gauging, measurement, and documentation examination.
11.1.2 Analysis
Verification by analysis will be used to show design margins. Also, when the particular tests
required for verification are impractical, risky, unacceptably long, or prohibitively expensive,
analysis may be used instead of testing, as noted in the verification matrices.
Analysis, including simulations where applicable, will also be used to guarantee that the Solar
Array Panels and their components will perform as expected under worst-case conditions.
11.1.3 Test
Verification by test includes, but is not limited to, the evaluation of performance by use of
special equipment or instrumentation, simulation techniques, and the application of established
principles and procedures to determine compliance with requirements.
11.2 INSPECTION REQUIREMENTS
Verification by inspection shall be by one of these three methods: 1) visual inspection of the
physical component; 2) a physical measurement of a property of the component, or; 3) a
documentation search demonstrating hardware of an identical design has demonstrated
fulfillment of a requirement.
11.2.1 Visual Inspection
Visual inspection of the physical component shall be performed by a customer appointed
qualified inspector to certify that the component has the properties/configuration specified in the
requirement.
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11.2.2 Physical Measurement
Physical measurement of component property (mass, dimensions, etc.) shall be performed by a
customer appointed qualified inspector to demonstrate the component meets specific
requirement.
11.2.3 Documentation Search
Verification of requirements based on similarity shall include supporting rationale and
documentation and be approved by the NASA/GSFC COR.
11.3 ANALYSIS REQUIREMENTS
Verification of performance or function through detailed analysis, using all applicable tools and
techniques, is acceptable with NASA/GSFC COR approval. Detailed descriptions of the
minimum required analyses, as well as analysis requirements, are provided in the SOW.
11.4 TEST REQUIREMENTS
a. This section provides general test requirements on how testing is to be performed in the
process of verifying that the deliverable item meets its requirements. Performance
parameter measurements shall be taken to establish a baseline that can be used to assure
that there are no data trends established in successive tests that indicate a degradation of
performance trend within specification limits that could result in unacceptable
performance in flight.
b. Any requirement that exceeds previous qualification test data shall be presented to the
NASA/GSFC COR as part of the verification planning process described in the SOW, for
evaluation and a possible delta qualification test.
11.4.1 Definitions
The hardware definitions are reproduced here from Section 1.8 of GEVS (GSFC-STD-7000A).
Acceptance Tests: The verification process that demonstrates that hardware is acceptable for
flight. It also serves as a quality control screen to detect deficiencies and, normally, to provide
the basis for delivery of an item under terms of a contract.
Design Qualification Tests: Tests intended to demonstrate that the test item will function within
performance specifications under simulated conditions more severe than those expected from
ground handling, launch, and orbital operations. Their purpose is to uncover deficiencies in
design and method of manufacture. They are not intended to exceed design safety margins or to
introduce unrealistic modes of failure. The design qualification tests may be to either “prototype”
or “protoflight” test levels.
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Protoflight Hardware: “Flight hardware of a new design; it is subject to a qualification test
program that combines elements of prototype and flight acceptance verification; that is, the
application of design qualification test levels and flight acceptance test durations.” The purpose
of the test on this hardware is to prove that a new design meets one or more of its design
requirements. Protoflight testing is performed at maximum expected flight levels plus a margin.
Test durations are typically the same as for acceptance tests.
Follow-On (Acceptance) Hardware: “Flight hardware built in accordance with a design that
has been qualified either as prototype or as protoflight hardware; follow-on hardware is subject
to a flight acceptance test program.” The purpose of the test on this hardware is to prove that a
particular flight unit has been manufactured properly. The design has already been proven
during a qualification or protoflight test program. Acceptance testing is performed at maximum
expected flight levels.
11.4.2 NA
11.4.3 Test Tolerances
Tolerances for the various mechanical test parameters are given in Error! Reference source not
found..
Table 11-1. Test Tolerances
Test Test Parameter Tolerance
Temperature 2 C
Humidity 5% RH
Mass Properties Weight: 25g
Center of Gravity: 6 mm
Moments of Inertia 10 %
Products of Inertia 10 %
Pressure >1.3 x 104 Pa (> 100 mm Hg): 5%
1.3 x 104 to 1.3 x 102 Pa (100 mm Hg to 1 mm Hg): 10%
1.3 x 102 to 1.3 x 101 Pa (1 mm Hg to 1 micron): 25%
< 1.3 x 101 Pa (< 1 micron): 80%
11.4.4 Test Restrictions
11.4.4.1 Failure During Tests
a. When a failure (non-conformance or trend indicating that an out-of-spec condition will
result) occurs, determination will be made as to the feasibility and value of continuing the
test to its specified conclusion. The test shall be stopped if equipment fails during testing
in cases where this failure will result in damage to the equipment.
b. Otherwise, the test shall be completed to obtain as much information as possible. If
corrective action is taken, the test will be repeated to the extent necessary to demonstrate
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that the test item’s performance is satisfactory. If corrective action taken as a result of
failure affects the validity of previously completed tests (e.g., redesign of a component),
prior tests will be repeated.
If during a test sequence, a test item is operated in excess of design life and wears out or
becomes unsuitable for further testing from causes other than deficiencies, a spare will be
substituted, and previously completed tests will be repeated to the extent necessary.
c. No replacement, adjustment, maintenance, or repairs are authorized during testing. This
requirement does not prevent the replacement or adjustment of equipment that has
exceeded its design operating life during tests, provided that after such replacement, the
equipment is tested as necessary to assure its proper operation. A complete record of any
exceptions taken to this requirement shall be included in the test report.
11.4.4.2 Modification of Hardware
Once the formal acceptance test has started, cleaning, adjustment, or modification of test
hardware shall not be permitted.
11.4.4.3 External Adjustment
The Solar Array Panels shall be designed so that no external adjustments are required after start
of acceptance or qualification testing.
11.4.4.4 Re-Test Requirements
If any event, including test failure, requires that a Solar Array Panel be disassembled and
reassembled, then all tests performed prior to the event shall be considered for repeat. If the unit
has multiple copies of the same build, then all units must be examined to determine if the
problem is common. If all copies require disassembly for repair, then each must receive the
same test sequence.
11.5 REQUIRED TESTS
The orders of tests on the insulated substrates, qualification coupons, and flight panels are
specified in the applicable sections below.
The Intermediate Sequence tests consist of the following functional tests and shall be performed
as specified during qualification coupon and flight panel testing:
Visual Inspection of Add-on Hardware (Everything the contractor has added to the
substrate) per Section 11.5.22 of this specification
Visual Inspection of Insulated Substrate
Adhesion Check
Test Condition Power Verification per Section 11.5.19 of this specification
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Substrate Insulation Resistance
Wiring Resistance
Resistance between each solar cell coverglass and the solar cell interconnect for the
applicable cell per Section 11.5.24 of this specification
Bypass Diode Functionality per Section 11.5.20 of this specification
11.5.1 NA
11.5.2 NA
11.5.3 NA
11.5.4 NA
11.5.5 NA
11.5.6 NA
11.5.7 NA
11.5.8 NA
11.5.9 NA
11.5.10 Thermal Vacuum Bake-out
The contractor shall bake out the Solar Array Panels prior to thermal vacuum cycling until the
outgassing requirements of Section 7.1.2.2 are met. The rate measured at the TQCM shall be
adjusted to account for chamber geometry, presence of cold sinks, chamber pumping speed, view
factors of the TQCM, and any other factors necessary to assure an accurate measurement of the
total outgassing per unit time per Kg mass of the unit under test. The chamber and GSE
containing organic materials shall be baked out and certified prior to installing the flight
hardware into the chamber per paragraph 11.5.10.3 below. NASA/GSFC shall perform the
calculations necessary to convert the outgassing rate to the test specific TQCM rate unless the
contractor elects to perform its own calculations and submits their assumptions, calculations and
results to NASA/GSFC for review. The cleanliness of the panels shall be preserved post bakeout
as described in paragraph 11.5.10.4 below.
The outgassing rate shall be considered acceptable when the flight hardware temperature does
not vary by more than 2 degrees C and the rate meets the requirements of Section 7.1.2.2.3 for 4
continuous hours.
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11.5.10.1 Test Conditions for Thermal Vacuum Bakeout
The contractor shall bake the panels at a temperature of 135+0/-5 °C in a vacuum of 1 x 10-5
Torr or less for at least 96 hours.
The contractor shall perform the out gassing verification at the end of the thermal vacuum
bakeout when flight hardware temperature is lowered to 97C, Panel Performance in Thermal
Vacuum Environment with no break in chamber vacuum.
The contractor shall maintain the TQCM at -20 C throughout the test to measure total volatile
out gassed condensables without the influence of water vapor. The TQCM must have a
representative view of the hardware.
11.5.10.2 Calculation of the Configuration Adjusted TQCM Rate
The contractor shall collect and deliver to the NASA/GSFC COR the information necessary to
model the test chamber. NASA/GSFC shall calculate the acceptance QCM rate that indicates
that the hardware has met the outgassing rate criteria. This rate shall be provided to the
contractor within 24 hours of the test starting providing that (1) any changes in the configuration
are minor and (2) they are provided to the GSFC contamination control analyst within 12 hours
of the test starting. When these conditions are not met the rates may take longer to supply. The
following information shall be provided to the NASA/GSFC COR at least one month prior to the
test and updated as necessary to account for unforeseen last minute changes in configuration:
Drawing or sketch showing chamber configuration and geometry
Chamber dimensions
Shrouds dimensions,
Temperature and physical dimensions of all major elements that will not be isothermal
(within 5 degrees) of the flight hardware (shrouds, heater plates, scavenger plates, LN2
lines, windows, etc.),
QCM location,
Cold finger, QCM heat sink, scavenger plate locations, sizes, control temperatures during
the test and general set-up
Pump location and dimensions
For the S-112 Qualification Coupon panel thermal vacuum setup, and with the QCM
temperatures for each bakeout in accordance with Section 11.5.10.1 of this specification, the
QCM goal for each bakeout shall be <635Hz/hr above background.
11.5.10.3 Pre-Test Chamber Vacuum Bakeout and Chamber Certification
Prior to the test with the flight hardware, the chamber shall be baked out in vacuum with any
GSE containing organic materials to establish the chamber background rate.
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The chamber shall be baked until the delta delta QCM rate is less than 5 Hz per hour with the
chamber at 130 degrees C. Once this is achieved, the chamber background rate shall be
measured for both certification temperatures specified in Section 11.5.10.1 with the chamber
conditions as specified in 11.5.10.1. Data shall be taken for a minimum of 4 hours at each
temperature. The chamber background rate shall be low and steady enough so as to not impede
measurement of the array outgassing rate. The PACE contamination analyst may relax these
requirements if later analysis proves that the measurement may be made.
Once this has been achieved, the contractor shall proceed to measure the background of the
chamber. Use of lower temperatures will lengthen the time required to complete the bakeout.
The following data shall be recorded during the test: QCM readings (once every 2 minutes
minimum), hardware temperature, chamber and shroud temperature, scavenger plate and other
cryosurface temperature, QCM temperature, and pressure.
If requested by the NASA/GSFC COR and prior to loading the chamber, the contractor shall
allow a minimum of two hours for a NASA/GSFC representative to inspect the chamber, its
equipment, and its configuration.
11.5.10.4 Post-Bakeout Cleanliness
The cleanliness of the flight hardware shall be maintained post bakeout. It is recommended that
the arrays be kept in sealed (i.e., low permeability) container or bagging material of the same
cleanliness level as the panels when the panels are in storage or during periods when a
contamination hazard is present in the area where the panels are located (painting, roof sealing,
curing of large quantities of polymers (adhesives/paints)).
11.5.10.4.1 Witness Foils
A NASA/GSFC provided aluminum witness foil shall be kept in the same environment (air
stream) as each panel at all times. The number of foils may be reduced to one if multiple panels
are kept in the same area or container. Each foil shall be replaced with a new foil and sent back
to the NASA/GSFC COR for analysis monthly. NASA/GSFC will send instructions on how to
handle and ship the foils with the foils. The collected non-volatile residue shall remain less than
0.2 mg/sq-ft per month. If this number is exceeded, the NASA/GSFC COR and the contractor
shall jointly determine a course of action. Normally deposited surface contaminants – for
example finger oils or phalates – should be able to be removed via a solvent cleaning of the
arrays. If there should be a contaminant that cannot be removed, then NASA/GSFC will conduct
further analysis to determine the volatility of the contaminant to determine if the bakeout shall
be required to be repeated.
Should a panel or group of panels need to be placed into a vacuum chamber after thermal
vacuum bakeout has been completed, the chamber shall meet the requirements of paragraphs
11.5.10.1. A QCM or witness foil that is kept at the same temperature as the flight hardware
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shall be installed into the chamber to verify that the cleanliness of the hardware has not been
compromised by the exposure to the chamber or test conditions.
NASA/GSFC shall supply a sufficient quantity of 18 inch by 12 inch aluminum witness foils to
the contractor to monitor the cleanliness of the air to which the panels are exposed. A handling
procedure shall be supplied with the foils. The foils shall be used to monitor cleanliness post
bake-out. If there is a concern for the facility cleanliness, they may also be used by the
contractor to monitor facility cleanliness. The foils shall be replaced monthly or if there is a
contamination event and sent back to NASA/GSFC for analysis.
11.5.10.5 Surface Cleanliness Required for Performance
The contractor shall supply the assumed surface cleanliness level used in calculations of EOL
solar array performance. It should be expressed in percent obscuration ratio per IEST-STD-
CC1246D.
11.5.11 Thermal Vacuum Test
11.5.11.1 Thermal Vacuum Test Parameters
The contractor shall thermal cycle each flight panel in a vacuum of 1 x 10-5 Torr or less. The
contractor shall perform the thermal vacuum cycling in a chamber with a shroud at
approximately liquid nitrogen temperatures.
The contractor shall cycle the flight panels from -95°C to +120°C for 9 cycles,
and from -100°C to +135°C for 3 cycles.
The contractor shall fix at least 6 calibrated temperature sensors over each flight panel and use
the panels’ platinum resistor thermometers to monitor temperature.
The contractor shall cycle to the temperature extremes based on the average reading of the
temperature sensors.
The temperature gradients across the panels during temperature transitions shall be limited to
±25C.
The period for one cycle shall be greater than 1.5 hour, excluding the dwell.
The dwell at the temperature extremes shall be greater than 1 hour.
The rate of temperature change between cold and hot limits shall not exceed 30°C per minute.
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The contractor shall send current through each panel circuit and through the platinum resistor
thermometers using the panel connectors during these tests.
The contractor shall continuously monitor the currents with a channel dedicated to that circuit,
heater, or resistor.
The contractor shall pass current through the cells during the first three cycles, pass current
through the bypass diodes during the second three cycles, pass current through the cells during
the seventh through ninth cycle, and pass current through the bypass diodes during the last three
cycles.
The reverse current through the cells shall be conducted through a flight quality test connector,
which the contractor shall mount to the flight panels that make connection to the anode side of
each string’s blocking diode and the service connector contacts that make connection to each
string’s return.
The Contractor shall not employ mechanical connections, other than flight quality crimps, during
the thermal vacuum exposure.
During the pump-down, the contractor shall monitor power line voltages, to demonstrate the
absence of corona discharge and multipaction.
Transitions from cold to hot conditions increase contamination hazards because material that has
accreted on the chamber walls may evaporate and deposit on the relatively cool solar panels.
Transitions will be conducted at rates sufficiently slow to prevent this from occurring. The
thermal vacuum test shall start and end with the flight hardware at 40 degrees C minimum to
minimize this risk and the thermal driving surfaces (shroud/heater) shall not be more than 40
degrees warmer than the flight hardware unless a QCM demonstrates that there is no risk.
Pass/Fail Criteria: Any discontinuity showing in recorded data shall fail the verification.
Any defects or reduction of power output outside the limits of the requirements of this
specification shall fail the verification.
Any mechanical damage to the panel, including the substrate and contractor installed
components, that put into question their ability to perform adequately during launch or flight
shall fail the verification.
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11.5.12 NA
11.5.13 NA
11.5.14 NA
11.5.15 NA
11.5.16 Solar Cell and Bypass Diode Qualification Tests
The Contractor shall use solar cells for PACE that have been qualified to either AIAA-S-111-
2005 or AIAA-S-111A-2014 and shall demonstrate previous flight heritage for the cells in a
space environment that envelopes that of the PACE mission.
11.5.17 Solar Array Panel Qualification Tests
a. The Contractor shall qualify the PACE solar array panels using AIAA S-112A-2013,
Qualification and Quality Requirements for Electrical Components on Space Solar Panels as
tailored by this Specification.
b. The contractor shall use one or more coupons for all of the following tests. Section
numbers below refer to S-112A unless otherwise indicated.
Life Cycles Coupon Test with Humidity Exposure, Section 7.1
Solar Absorptance Characterization, Section 8.4
Bypass Diode Characterization, Section 8.5
The following tests are waived:
Humidity Exposure, Section 7.1.3.1
Acoustic Exposure, Section 7.2.4.
ESD Test, Section 7.3
UV Radiation Effects Characterization, Section 8.1
Angle of Incidence Characterization, Section 8.2
Atomic Oxygen Characterization, Section 8.6
c. The following tests or requirements are modified:
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Sample Count, Section 6.1: The minimum sample count for CICs on the life cycle coupons shall
be the maximum number that will fit on the front sides of the GFE qualification coupon
substrates.
Life-Cycle Coupons/Subcoupons, Section 7.1.2: add the requirement that three cells be broken
and repaired prior to the start of environmental exposure. The contractor shall break one cell at a
positive string termination, another cell at a negative string termination, and another in the center
of a row.
Combined Effects Exposure, Section 7.1.3.4, Subsection 4 of S-112A: Dark forward bias may be
substituted for illumination during the heated portion of the cycles.
The VCM test in Section 7.2 shall be performed on a Life Cycle Coupon.
11.5.17.1 Solar Array Panel Life Cycle Coupon Tests
Section 7.1 of S-112A requires life cycle coupons. The coupons used in the cycling test must
include at least two samples of all of the parts to be used on the solar array panels.
a. In S-112A, the bakeout temperature shall be 135°C+0/-3°C.
b. Except for the 135°C hot survival temperature and the -100° C cold survival temperature,
the temperature extremes for the life cycle coupon test shall be 10°C beyond the
extremes and for 1.5X the number of cycles in Error! Reference source not found. of
this specification. The cycles to the hottest temperature shall not exceed 135°C.
c. The order of tests on the life cycle coupon shall be as follows:
Intermediate Sequence as specified in Section 11.5:
Damage and Repair
Panel Mass Measurement
Intermediate Sequence
Thermal Vacuum Bakeout/VCM Test per Sections 11.5.17and 11.5.17.1 of this specification
Thermal Vacuum Cycles 1 through 8
Intermediate Sequence
Rapid Temperature Cycles 1 through 10
Intermediate Sequence
Rapid Temperature Cycles 11 through 100
Intermediate Sequence
Rapid Temperature Cycles 101 through 5,000
Intermediate Sequence
Rapid Temperature Cycles 5,001 through 10,000
Intermediate Sequence
Rapid Temperature Cycles 10,001 through 24,074 (1.5X number of mission cycles)
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Intermediate Sequence
d. Pass/Fail Criteria: The coupons shall fail the thermal cycle test if the test condition power
degrades by more than 3% from the baseline measurement prior to the start of cycling.
11.5.18 Flight Solar Array Panel Tests
The order of tests on the flight panels shall be as follows:
Intermediate Sequence as defined in Section 11.5 of this specification
Thermal Vacuum Cycles per Section 11.5.11 of this specification.
Thermal Vacuum Bakeout per Section 11.5.10 of this specification.
Intermediate Sequence
Panel Mass Properties Measurement per PACE-PWR-SOW-0025, PACE Solar Array Panels
Statement of Work.
11.5.19 Test Condition Power Verification
a. The contractor shall use a solar simulator calibrated with a set of primary or secondary
standard solar cells to determine the current-voltage (I-V) characteristics of each panel
through the panel terminals at 23º ± 5ºC.
b. The contractor shall extrapolate the measured data to obtain the I-V curve for a panel
operating at 28ºC and AM0.
Type: Measurement.
Level: Panel.
Schedule: Per Sections 11.5.17, 11.5.18, and 11.5.19
c. Pass/Fail Criteria: The verification shall fail if the power at the specified load voltage of
a panel does not meet the requirements of Section 4.1.1.
11.5.20 Bypass Diode Functionality Verification
a. The contractor shall pass a minimum of 120% of the Isc of each string on each panel
through the bypass diode circuits. This test shall be conducted at 23º ± 5ºC with the
results extrapolated to 28ºC.
Type: Test.
Level: Panel.
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Schedule: Per Sections 11.5.17 and 11.5.18.
b. Pass/Fail Criteria: The voltage dropped by the bypass diodes shall not vary more than
±3% from the first panel level measurement to the last at an extrapolated temperature of
28ºC.
11.5.21 Substrate Insulation Resistance Verification
a. The contractor shall make connection to the cell circuits through each panel’s terminals
and shall tie all positive and negative power leads together.
b. The contractor shall then measure the resistance between these tied together leads and the
panel’s substrate.
c. The contractor shall make this measurement at 500 volts direct current (Vdc) with the
current limited to 20 microamperes or less with the positive test voltage on the cell
circuits
d. Type: Measurement. The insulation resistance values shall be recorded.
Level: Panel.
Schedule: Per Sections 11.5.17 and 11.5.18
e. Pass/Fail Criteria: The panel shall fail verification if its substrate insulation resistance is
less than that required by Section 4.1.11.1.
11.5.22 Solar Cell Mechanical Verification
a. The contractor shall visually inspect each solar cell on the qualification and flight panels
for compliance with Section 4.1.5.1.
b. The contractor shall inspect with the unaided eye and under a minimum of seven power
magnification.
c. The contractor shall perform optional inspections, which it determines are
advisable, at its discretion.
Type: Inspection.
Level: Panel.
Schedule: Per Sections 11.5.17 and 11.5.18.
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d. Pass/Fail Criteria: Each solar cell shall meet the requirements of Section 4.1.5.1 or the
contractor shall remove the cell from the panel and replace it.
e. If more than 4% of the cells on the cell qualification coupon shall crack as a result of test,
the panel shall fail qualification.
11.5.23 Cover Orientation Verification
a. The contractor shall visually inspect each solar cell cover for compliance with Section
4.1.6.2.
Type: Inspection.
Level: Panel.
Schedule: Per Sections 11.5.17 and 11.5.18.
b. Pass/Fail Criteria: No more than 0.1% of a panel’s covers shall fail this test.
11.5.24 Cover Grounding Verification
a. The contractor shall verify the electrical connection between the geometric center of each
cover within ± .5 cm, and its connection to ground or the array circuitry at 23º ± 5ºC
b. Pass/Fail Criteria: The electrical connection shall fail the test if the resistance exceeds 1 x
109 ohms.
11.5.25 Flight Connector Type Verification
The contractor shall propose the type, level, schedule and pass/fail criteria for verifying the
connector type, connector mounting adequacy, and wiring.
11.5.26 Platinum Resistance Thermometer Type Verification
The contractor shall propose the type level, schedule and pass/fail criteria for verifying the
resistor type.
11.5.27 Platinum Resistance Thermometer Performance Verification
Type: Measurement.
Level: Component.
Schedule: Per Sections 11.5.17 and 11.5.18.
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a. Pass/Fail Criteria: The output from each platinum resistor thermometer shall meet the
requirements of 4.1.10 or this requirement is failed.
b. During the thermal vacuum test or thermal cycling test, the platinum resistor thermometer
shall additionally be shown to operate without discontinuity in resistance or this
requirement is failed.
11.5.28 Parts and Assembly Layout Verification
The contractor shall propose methods and schedules for verifying the parts and assembly layout.
11.5.29 Mission Life Verification
a. The contractor shall certify that it has conducted a test program to demonstrate that the
solar panels have the required mission life.
b. The contractor shall certify that it has not knowingly limited the mission life of the
panels to less than the required mission life.
11.5.30 Shelf Life Verification
The contractor shall certify that the solar panels will have a shelf life greater than 5 years when
packaged to its specifications.
11.5.31 Substrate Ground Verification
The contractor shall verify the substrate ground with an ohm meter.
11.5.32 Cleanliness Verification
a. Immediately prior to shipment, the contractor shall inspect the panels to the requirements
of JSC-SN-C-005 to the highly sensitive level using both white and black light
inspections.
b. Immediately prior to shipment, the contractor shall verify the cleanliness of the array by
blacklight inspection per IEST-STD-CC1246D.
Type: Inspection.
Level: Panel.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
Effective Date: December 18, 2018
61 Check https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use
Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.
400-FORM-0002 (4/16/2014)
Schedule: Prior to delivery to NASA/GSFC.
c. Pass/Fail Criteria: The inspection shall meet the visibly clean highly sensitive
requirements of JSC-SN-C-005.
d. If not, the contractor shall clean the “dirty” areas until the requirement is met.
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
Effective Date: December 18, 2018
62 Check https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use
Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.
400-FORM-0002 (4/16/2014)
APPENDIX A ABBREVIATIONS AND ACRONYMS
Abbreviation/
Acronym Definition
Al Aluminum
AM0 Air Mass Zero
AR/ITO Antireflective/Indium Tin Oxide Coating
CCB Configuration Control Board
CCR Configuration Change Request
CG Center of Gravity
CM Configuration Management
CMO Configuration Management Office
COR Contracting Officer's Representative
CSS Coarse Sun Sensor
CVCM Collected Volatile Condensable Mass
DC Direct Current
DDD Displacement Damage Dose
DILS Deliverable Items List and Schedule
EEE Electrical, Electronic, and Electromechanical
EOL End of Life
ESD Electrostatic Discharge
FS Factor of Safety
GFE Government Furnished Equipment
GSE Ground Support Equipment
GSFC Goddard Space Flight Center
ICD Interface Control Document
I&T Integration and Test
ITO Indium Tin Oxide
KOZ Keep-Out Zone
MLI Multi-Layer Insulation
Mohms Megaohms
MS Margin of Safety
NASA National Aeronautics and Space Administration
NIEL Non-Ionizing Energy Loss
PACE Plankton, Aerosol, Cloud, ocean Ecosystem
QCM Quartz Crystal Monitor
S/C Spacecraft
SOW Statement of Work
TBD To Be Defined
TBR To Be Reviewed
TID Total Ionizing Dose
PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A
Effective Date: December 18, 2018
63 Check https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use
Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.
400-FORM-0002 (4/16/2014)
Abbreviation/
Acronym Definition
TML Total Mass Loss
UUT Unit Under Test
VDA Vapor Deposited Aluminum
VDC Voltage, Direct Current
VCM Volatile, Condensible Material