Attachment K - L10 Space Environmental Req Document (LSERD).pdf

Attachment No. K

Landsat 10 Space Environmental Requirements Document (LSERD)

RFP 80GSFC26R0015

National Aeronautics and Space Administration Goddard Space Flight Center

Landsat 10 LSERD L10-SYS-REQ-0009 Revision B ii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Landsat 10 Space Environment Requirements (LSERD) Signature/Approval Page

Prepared by:

Electronic Signature in TDMS Approval Date in TDMS Michael Pryzby Date Landsat 10 Space Environments Engineer NASA/GSFC, Code 426

Approved by:

Cagatay Aymergen Date Landsat 10 Deputy Project Manager / Acting Observatory Manager

Ross Henry Date Landsat 10 Project Manager

Evan Webb Date Landsat 10 Mission Systems Manager NASA/GSFC, Code 533 iii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

CM Foreword This document is a Landsat 10 Project Configuration Management (CM)-controlled document.

Note: Prior to February 2026, the project name was Landsat Next (LNext).

For the purposes of configuration management, the prefixes “LNEXT” and “L10” are completely interchangeable. For example, LNEXT-MGMT-PLAN-0004 is the same as L10-MGMT-PLAN- 0004.

Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the Landsat 10 CM Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

Questions or comments concerning this document should be addressed to:

NASA/Goddard Space Flight Center Landsat 10 Project Office, Code 426 Attention: Configuration Management Office Greenbelt, Maryland 20771 iv Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

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

Table of Contents SIGNATURE/APPROVAL PAGE ................................................................................................ II CM FOREWORD ..........................................................................................................................III CHANGE HISTORY LOG .......................................................................................................... IV

TABLE OF CONTENTS ................................................................................................................ V

LIST OF FIGURES ...................................................................................................................... VI LIST OF TABLES ....................................................................................................................... VII

1.0 INTRODUCTION

1.1 Purpose

1.2 Scope

1.2.1 Identification

1.2.2 Applicability

1.2.3 Terminology

1.2.4 Document Overview

1.2.5 Rationale and Notes

1.3 Applicable and Reference Documents

1.3.1 Applicable Documents

1.3.2 Project Applicable Documents

1.3.3 Landsat 10 Reference Documents

1.3.4 Project Reference Documents

2.0 MISSION OVERVIEW

2.1 Mission Statement

2.2 Mission Background

2.3 Mission Objectives

2.4 Mission Implementation

2.4.1 Space Segment

2.4.2 Ground Segment

2.4.3 Launch Segment

3.0 ENVIRONMENTAL CONDITIONS

3.1 Charged Particle Radiation Environment

3.1.1 Total Ionizing Dose (TID) Environment

3.1.1.1 TID Level Requirements

3.1.2 Total Non-Ionizing Dose / Displacement Damage Dose Environment (DDD) .10

3.1.2.1 DDD Level Requirements

3.1.3 TID/DDD Testing

3.1.4 Spot Shielding

3.1.5 Single-Event Effects

3.1.5.1 Non-Destructive SEE (e.g., SEUs, SETs, SEFIs, SHEs)

3.1.5.2 Destructive SEE

3.2 Micrometeoroid and Orbital Debris (MMOD) Environments

3.3 Observatory Magnetic Fields

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

3.4 Atomic Oxygen

3.5 Observatory Charging from All Sources

3.5.1 Observatory Charging Potential

3.5.2 Observatory Charging Electric Field

3.5.3 Observatory Grounding to Chassis Ground Plane (CGP)

3.5.3.1 Grounding of Dielectric Films and Bulk Materials

3.5.3.2 Thermal Blankets

3.5.4 Solar Array Surface Charging

3.5.5 Sensor Shielding

APPENDIX A PLASMA ENVIRONMENT SPECIFICATION

Space Charging Environment Surface Charging Environment Model Parameters Bulk Charging (Energetic Particle) Environment

APPENDIX B REQUIREMENT APPLICABILITY MATRIX

APPENDIX C MICROMETEOROID ORBITAL DEBRIS PROTECTION TO IMPROVE THE

RETURN OF MISSION SCIENCE

APPENDIX D ABBREVIATIONS AND ACRONYMS

APPENDIX E RADIATION EFFECTS DEFINITIONS

Cumulative Effects Instantaneous Effects Energy deposition

List of Figures Figure 2.4-1 Landsat 10 Mission Architecture Figure 3.2-1 Design Curves for Protection against Penetration of 1.4mm 14.5 km/s particles Figure 3.4-1 Atomic oxygen flux (atoms/cm2 /month) for 63-month (5- year mission plus

3-month commission with accumulated AO fluence for + 2σ solar activities Figure A-1 Mission Average Integral Omnidirectional Current Density Projected Through a Flat Surface, (nA/cm2) Figure A-2 Worst-Case Omnidirectional Integral Flux Projected Through a Flat Surface

(nA/cm2) Figure C-1 Overall Description of the Proposed Process Figure C-2 Flux Variations Across Years for 650 km Vs 705 km Orbits Figure C-3 Orbital Debris Flux by Particle Diameter for both 650 km and 705 km Orbits

Figure C-4 Omnidirectional Orbital Debris Flux Curve for the Landsat 10 Orbit Figure C-5 Directionality of the Orbital Debris Threat to Landsat 10 Figure C-6 Design Curves for Protection against Penetration of 1.8mm 14.7 km/s particles

Figure C-7 Design Curves for Protection against Penetration of 1.4mm 7.25km/s particles vii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

List of Tables Table 2.4.1-1 Landsat 10 Mission Parameters Table 3.2-1 Blanket Areal Density (g/cm2) at Given Wall Thickness, 14.5 km/s Table 3.5-1 Resistivity Definitions Table 3.5.3.2-1 Table Ground Bond Connections per Blanket Area Table A-1 Parameters for Surface Charging Table A-2 Mission Average Integral Omnidirectional Current Density Projected Through a Flat Surface, nA/cm2 Table A-3 Mission Average Integral, Omnidirectional Number Flux (Electrons/cm2-sec )

Table A-4 Worst-Case Omnidirectional Integral Flux Projected Through a Flat Surface

(nA/cm2) Table A-5 Worst-Case Integral Omnidirectional Flux (Electrons/cm2/s) Table C-1 Directional Orbital Debris Flux (#/m2-yr) and Velocity (km/s) for the Landsat

10 Orbit Table C-2 Numerical Values for the Data Plotted in Figure C-6 Table C-3 Numerical Values for the Data Plotted in Figure C-7, for 7.25 km/s case Table C-4 Increase in Effective Wall Thickness from Layers of Kevlar

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

1.0 INTRODUCTION

1.1 PURPOSE

The Landsat 10 Space Environments Requirements Document (LSERD) establishes a set of requirements for safe operations of the Landsat 10 Observatory (spacecraft and instruments) in their predicted orbit and space environment.

1.2 SCOPE

1.2.1 Identification

This Landsat 10 (Landsat 10) Space Environments Requirements Document (LSERD) sets forth the general space environment for the Landsat 10 mission. The environments addressed include radiation, space charging, magnetics, atomic oxygen, and micrometeorite orbital debris flux.

1.2.2 Applicability

This document covers the Landsat 10 spacecraft and Landsat instrument suite (LandIS).

1.2.3 Terminology

Throughout this document the following definitions are used:

The term "spacecraft" is interchangeable with spacecraft bus or bus, and it refers to all parts of the space segment that are not the instruments.

The term "observatory" refers to the fully or partially integrated system, including the spacecraft and LandIS. This term is used mainly in reference to testing in order to differentiate between the spacecraft bus without the instrument and the integrated satellite with instrument.

The expression “instrument provider” refers to the organization / company delivering the instrument for integration on the spacecraft. The term is used interchangeably with the instrument contractor or sensor subcontractor with respect to Landsat 10.

Additional radiation effect definitions are defined in Appendix E.

1.2.4 Document Overview

This LSERD is controlled and maintained by the Landsat 10 Project office. The LSERD is imposed on the instrument contractor(s) and the spacecraft contractor and will serve as a starting point for a detailed set of requirements defining the on-orbit environment.

1.2.5 Rationale and Notes

Throughout this document, there are rationale and note sections which are provided for information or clarification, but do not represent requirements. With such complimentary text, it is possible that the same or derived information may exist as a requirement, but any requirement is properly numbered.

In general, exceptions to all limits and levels can, if applicable, be made with a probability and confidence interval such that the intent of this document is satisfied. This is particularly true for multiplicative factors of margin on design and explicit Linear Energy Transfer (LET) thresholds.

Exceptions to requirements are reviewed and approved by the Parts Control Board (PCB). In the event that requirements cannot be met through analysis, only then will a waiver process be initiated.

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

Where flight lot testing is called for or applicable, requirements associated with mission reliability will determine sufficient sampling for lot testing. Flight lot requirements may only be removed with sufficient sampling, representative lot-to-lot data and application verification with the system, or design engineer (required from the onset of design), subject to PCB approval.

1.3 APPLICABLE AND REFERENCE DOCUMENTS

1.3.1 Applicable Documents

The LSERD is consistent with and responsive to the following revisions of the applicable documents. These documents establish detailed specifications, requirements, and interface information necessary for the performance of the contract. The contactor will comply with the applicable documents below as they apply to the performance of their Landsat 10 contract.

Unless otherwise stated in this document, all inconsistencies in the LSERD will be resolved as defined in the contractor's Statements of Work.

Document Number Title Revision NASA-STD-4003 Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads, and Flight Equipment A w/CHANGE 1

NASA-STD-8719.14 Process for Limiting Orbital Debris C NPR 8715.6 NASA Procedural Requirements for Orbital

Debris Mitigation E

In this document, citations are assumed to be the latest version unless otherwise noted.

This document table was generated from the Landsat 10 Common Reference and Applicable Documents List (L10-MGMT-DESC-0008), DRAFT Rev F.

1.3.2 Project Applicable Documents

Document Number Title L10-SYS-DESC-0015 Landsat 10 Radiation Environment Description (LRED) In this document, citations are assumed to be the latest version unless otherwise noted.

1.3.3 Landsat 10 Reference Documents

The LSERD is consistent with the following documents. Unless otherwise stated in this document, all inconsistencies in the LSERD will be resolved as defined in the contractor’s Landsat 10 Spacecraft Statement of Work (SC SOW), L10-SC-SOW-0004 and Landsat 10 Landsat Instrument Suite (LandIS) Statement of Work (SOW), L10-LANDIS-SOW-0003.

Document Number Title Revision NASA Alert

NA-GSFC-2005-05

NASA Advisory, Latent Damage Reliability Issues Due to Single-Event Latchup and High-Current Error Modes

N/A

NASA-HDBK-4002 Mitigating In-Space Charging Effects – A Guideline B w/CHANGE 1 NASA-HDBK-4006 Low Earth Orbit Spacecraft Charging Design Handbook A NASA-STD-4005 Low Earth Orbit Spacecraft Charging Design Standard A w/Change 1 NASA/TM-2018-220074 Guidelines for Verification Strategies to Minimize RISK

Based on Mission Environment, -Application and - Lifetime (MEAL), O. Gonzalez, Y. Chen, R.L. Ladbury, D.R. Morgan, C.M. Green and D.E. Yuchnovicz

N/A

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

Document Number Title Revision NASA/TM-2019-220269 Radiation Single Event Effects (SEE) Impact on

Complex Avionics Architecture Reliability, R.F. Hodson, D. Morgan, R.L. Ladbury, Y. Chen, M. Bay and J.

Zinchuk

N/A

SEE Guideline Proton Single Event Effects (SEE) Guideline, NASA Electronics Parts and Package (NEPP) Program

N/A

Single Event Effect Criticality Analysis (SEECA) https://radhome.gsfc.nasa.gov/radhome/papers/seecai.htm N/A

AFRL-RV-PSTM-2023-0003 NASA/Air Force Spacecraft Charging Analyzer Program (NASCAP-2K) Version 4.5 User's Manual

N/A

MFS-32056-1 NASA/Air Force Spacecraft Charging Analyzer Program (NASCAP-2K), Version 4.5

N/A

ECSS-E-ST-20-06C Spacecraft charging (15 May 2019) 1 2015 IEEE NSREC Short Course

Design Challenges for Optical Payloads Used in the Space Radiation Environment

N/A

2017 IEEE NSREC Short Course

Strategies for SEE Hardness Assurance—From Buy-It- And- Fly-It to Bullet Proof

N/A

2017 IEEE NSREC Short Course

Total Ionizing and Non-Ionizing Dose Radiation Hardness Assurance

N/A

IEEE Transactions on Nuclear Science, Vol. 40, No. 6, pp.

1372-1379, 1993

Damage Correlations in Semiconductors Exposed to Gamma, Electron, and Proton Radiation, Summers, G.P., et al.

N/A

IEEE Transactions on Nuclear Science, Vol. 50, No. 6, pp.

2341-2351, 2003

SEE Characterization of Vertical DMOSFETs:

An Updated Test Protocol, Titus, J.L., and Wheatley, C.F.

N/A

IEEE Transactions on Nuclear Science, Vol. 55, No. 6, pp 3122-3129, 2008

Recommended Test Conditions for SEB Evaluation of Planar Power DMOSFETs, Liu, S., et al.

N/A

IEEE Transactions on Nuclear Science, Vol. 56, No. 6, pp.

3354-3360, 2009

Statistical Model Selection for TID Hardness Assurance, Ladbury, R., et al.

N/A

IEEE Transactions on Nuclear Science, Vol. 64, No. 1, pp.

325-331, 2017

Inclusion of Radiation Environment Variability in Total Dose Hardness Assurance Methodology, Xapsos, M.A., et al.

N/A

Journal of Spacecraft and Rockets, Volume 50, Number 6, pp 1277-1287

Test Approach to Assess Solar Array Diode Board Susceptibility to Sustained Arcing, M.J. Bodeau, doi:

10.2514/1.A32442

N/A

MIL-STD-883 Department of Defense, Test Method Standard, Microcircuits

L

MIL-STD-883-3 Department of Defense, Test Method Standard, Environmental Test Methods for Microcircuits, Part 3:

Test Methods 3000-3999

L w/ Change 1

MSC-26690-1 Debris Assessment Software (DAS) 3.2.6 2012 Space Power Workshop Sustained Arc Qualification Test of an Array Diode

Board, M.J. Bodeau N/A

In this document, citations are assumed to be the latest version unless otherwise noted.

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

1.3.4 Project Reference Documents

Document Number Title L10-LANDIS-ICD-0001 Landsat Instrument Suite (LandIS) Interface Requirements Document

(LIRD)

L10-LANDIS-SOW-0003 Landsat Instrument Suite (LandIS) Statement of Work (SOW) L10-SC-SOW-0006 Landsat 10 Spacecraft Statement of Work (SC SOW) In this document, citations are assumed to be the latest version unless otherwise noted.

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

2.0 MISSION OVERVIEW

2.1 MISSION STATEMENT

Landsat 10, consistent with United States (U.S.) law and government policy, will continue the Landsat program's acquisition, archival, and distribution of multi-spectral imagery, thereby affording global, synoptic, and repetitive coverage of the Earth's land surfaces and coastal regions at a scale where natural and human-induced changes can be detected, differentiated, characterized, and monitored over time.

2.2 MISSION BACKGROUND

Following the successful launch of Landsat 8 (formally named Landsat Data Continuity Mission, LDCM) in February 2013, and during the development of Landsat 9, the United States Geological Survey (USGS) and the National Aeronautics and Space Administration (NASA) recognized the need to assemble a team of experts from within both agencies to evaluate and inform an acquisition strategy for the Landsat mission to follow Landsat 9. The NASA-USGS Joint Agency Sustainable Land Imaging (SLI) Architecture Study Team (AST) was formed in September 2018, and tasked with investigating how to best satisfy the diverse set of user needs collected in the USGS User Needs for the Sustainable Land Imaging Program July 2018, Release

2.0. This investigation resulted in a set of recommendations to the headquarters of both agencies, delivered in December 2019. The highest-recommended “Roadmap 1” architecture described a small constellation of “superspectral” space-based sensors that would substantially improve the spectral, spatial, and temporal capabilities of previous Landsat missions, while continuing to satisfy the primary goal of ensuring a highly calibrated data set that maintains compatibility with the legacy data of the Earth’s land surfaces and coastal regions held in the National Satellite Land Remote Sensing Data Archive at the USGS’s Earth Resources and Observation Science (EROS) Center. In April 2020 NASA/Goddard Space Flight Center (GSFC) received authorization to initiate the Landsat Next project consistent with the AST’s “Roadmap 1” recommendation. Landsat Next subsequently passed its Mission Concept Review (MCR) and Key Decision Point-A (KDP-A) reviews in 2022. In March 2026, the Landsat Next mission was re-architected and re-designated Landsat 10 as a single observatory mission.

The goal of Landsat 10 is to continue the acquisition, archival, and distribution of multi-spectral imagery, thereby affording global, synoptic, and repetitive coverage of the Earth's land surfaces and coastal regions at a scale where natural and human-induced changes can be detected, differentiated, characterized, and monitored over time. This goal is in keeping with the Landsat programmatic goals stated in both the Commercial Space Act of 1998, Public Law #105-303 and the Land Remote Sensing Policy Act of 1992, Public Law #102-555. This policy requires that the Landsat Program provide data into the future that is sufficiently consistent with previous Landsat data to allow the detection and quantitative characterization of changes in or on the land surface of the globe.

Landsat 10 continues the long-running partnership of NASA and the USGS, with NASA providing the space and launch segments and the USGS providing the ground system (GS), providing the longest continuous global record of the Earth’s land surface and coastal regions.

The Landsat series of satellites have continuously acquired multispectral images of the global land surface since the launch of the Earth Resources Technology Satellite (ERTS, later renamed Landsat 1) in 1972.

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2.3 MISSION OBJECTIVES

Landsat 10 has these major mission objectives:

· Collect and archive moderate resolution Visible to Shortwave-Infrared (VSWIR) reflective and Thermal Infrared (TIR) emissive image data affording seasonal coverage of the global landmass for a continuous period of not less than five (5) years.

· Ensure that Landsat 10 data are sufficiently consistent with data from the earlier Landsat missions in terms of relative acquisition geometry, calibration, coverage characteristics, spectral characteristics, output product quality, and data availability; all of which permit scientific studies, management of land cover and land use change, along with associated inland and coastal water resources, over multi-decadal periods.

· Ensure Landsat 10 is responsive to documented emerging user needs and applications as characterized by periodic assessment, currently the User Needs for the SLI Program July 2018, Release 2.0, and identified by the operational requirements for acquisition, processing, archiving, and distribution of Earth’s image data to the United States Government and other users.

· Distribute Landsat 10 data products to the public freely on a nondiscriminatory basis.

2.4 MISSION IMPLEMENTATION

The Landsat 10 mission includes a single observatory in a 653 km sun-synchronous low-earth orbit (LEO). The observatory includes a single integrated suite of sensors called LandIS (the Landsat Instrument Suite) that images the full swath required to achieve global coverage via a repeating ground track every 18 days.

The LandIS improves all spectral and spatial capabilities by approximately twofold when compared with the Operational Land Imager (OLI) and Thermal InfraRed Sensor (TIRS) instruments flown on Landsats 8 and 9, while maintaining the historical Landsat data continuity and similar temporal coverage.

Each Landsat 10 partner agency has specific responsibilities and will deliver major elements to the overall mission. NASA will provide the Space Segment (SS) and the Launch Services Segment (LSS), and the USGS will provide the Ground System (GS) and Mission Operations.

NASA/Goddard Space Flight Center (GSFC) will provide overall Landsat 10 project management, mission system engineering and mission assurance during development and will transition the mission to USGS for science operations following on-orbit commissioning.

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Figure 2.4-1 Landsat 10 Mission Architecture

(Generated from Landsat 10 Common Mission Architecture Diagram L10-SYS-DESC-0022 Rev B)

2.4.1 Space Segment

NASA/GSFC provides the Space Segment via competitive procurements for the science instrument suite (LandIS) and the spacecraft bus. The Space Segment will consist of a single observatory flying in a sun-synchronous, near-polar orbit at 653 km altitude. The spacecraft, instrument, and overall Landsat 10 mission are specified as risk class C per NPR 8705.4. The overall mission parameters needed to achieve global coverage are shown below in the Mission Parameters Table.

Table 2.4.1-1 Landsat 10 Mission Parameters

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LandIS is a whiskbroom scanner which provides 26 spectral bands covering a swath on the ground of approximately 164 km on a fixed ground track defined by a world-wide reference system known as WRS-3. LandIS acquires raw detector data for all spectral bands within 15 seconds to minimize band-to-band and georegistration errors, reduce classification change detection errors due to clouds, and provide simultaneous acquisitions for science applications.

LandIS then performs data compression prior to packaging the data for transfer to the spacecraft storage and subsequent downlink. LandIS also provides internal calibration sources such as blackbody, a space port, stim lamps and solar diffuser.

The Space Segment will also provide simulators for mission, ground system and interface testing during the development phase of the mission and in support of long-term maintenance of systems. These simulators will have varying fidelity and capabilities to meet various mission needs.

2.4.2 Ground Segment

The Landsat Ground Segment (GS) consists of the ground system and its external interfaces, including NASA institutional services. The GS includes the Mission Operations Center (MOC), the Ground Network (GN), and the Data Processing and Archive System (DPAS). External interfaces include NASA's Near Space Network (NSN), GSFC Conjunction Assessment and Risk Analysis (CARA) and its Flight Dynamics Facility, along with other external interfaces.

The Landsat 10 MOC will be used for command and control, health and status monitoring, orbit and attitude maintenance, performance analysis, onboard memory management, and flight software maintenance. The MOC will support onboard autonomy for mission planning and scheduling. The MOC ingests, processes and archives telemetry data delivered via the GN. The Landsat 10 Flight Operations Team (FOT) located at the MOC performs these functions and is responsible for monitoring the observatory, detecting, investigating and resolving spacecraft anomalies and monitoring the instrument image collections and special image collections.

The GN provides geographically dispersed ground station resources for mission execution; it includes the Landsat Ground Network (LGN) and a wideband data routing capability to the Landsat cloud. The LGN provides communication capability to and from the observatory and MOC for commanding and housekeeping data via the S-band, and will receive science mission data from the observatory over a Ka-band downlink.

The DPAS ingests, processes, and archives science data returned from the observatory via the GN. LandIS data will be processed into data products with a ground sample distance (GSD) ranging from 10m-60m depending on the particular spectral band. The DPAS also provides a long-term archive capability for raw data and allows the user community to query, download, and directly interact in the cloud with Landsat 10 data products, via a public-facing web portal for receiving data products.

The USGS will lead overall Landsat 10 GS development. The USGS will also lead integration of the GS and ensure timely completion of ground readiness testing in preparation for NASA-led mission readiness activities. The MOC will support Landsat 10 mission readiness activities.

2.4.3 Launch Segment

The Launch Segment will provide the assets and services associated with the Launch Vehicle (LV) and the observatory-to-LV integration. The Launch Segment will include the LV; all

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Launch Vehicle-Ground Support Equipment (LV-GSE), property, and facilities to integrate the observatory to the LV and verify its integration; and perform prelaunch testing with ground-based functions. The launch vehicle and launch site will be specified during late Mission Phase B or early Phase C.

Sections 2.0-2.4.3 were generated using Landsat 10 Common Boilerplate (L10-MGMT-DESC- 0005) Rev E

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3.0 ENVIRONMENTAL CONDITIONS

This section specifies the environmental characteristics in the presence of which the spacecraft and the instrument components must meet all other requirements.

3.1 CHARGED PARTICLE RADIATION ENVIRONMENT

The Landsat 10 mission orbit radiation environment will consist of protons and electrons trapped in the Van Allen radiation belts; protons and heavier ions from solar events; Galactic Cosmic Ray ions; and solar wind plasma consisting of low-energy electrons, protons, and heavier ions.

3.1.1 Total Ionizing Dose (TID) Environment

LSERD-121 Effects of Total Ionizing Dose (TID) shall not cause permanent damage to, or performance degradation of, the flight system.

3.1.1.1 TID Level Requirements

LSERD-124 All Electrical, Electronic, Electromechanical, and Electro-optical (EEEE) parts shall function after exposure to the 95th percentile mission total ionizing dose as specified in the Landsat 10 Radiation Environment Description (LRED), L10-SYS-DESC-0015, Figure 5.2.1-1.

3.1.2 Total Non-Ionizing Dose / Displacement Damage Dose Environment (DDD) LSERD-146 Effects due to Displacement Damage Dose (DDD) shall not cause permanent damage to, or performance degradation of the flight system.

3.1.2.1 DDD Level Requirements

LSERD-149 All EEEE parts shall function after exposure to the 95th percentile mission displacement damage dose, as calculated from the Landsat 10 Radiation Environment Description (LRED), L10-SYS-DESC-0015 Figures 5.3.1-1 and 5.3.1-2 for the Landsat 10 mission for the appropriate material (most commonly silicon or gallium arsenide).

3.1.3 TID/DDD Testing

Flight lot testing is recommended if part test data (same part number, same manufacturer, and same process) does not exist for technology types susceptible to TID or DDD degradation.

3.1.4 Spot Shielding

If necessary, appropriate spot shielding (including Rad-PAK) may be added around a component so that the part can meet the shielded dose.

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The dose for the amount of shielding can be determined either by solid angle sectoring/three-dimensional ray tracing coupled with radiation transport simulation or by the dose versus spherical shielding depth values and can be calculated/modeled for specific amounts of shielding not found in the tables at the 95% confidence level in the Landsat 10 Radiation Environment Description (LRED), L10-SYS-DESC-0015, Figures 5.2.1-2, 5.3.1-1, and 5.3.1-2.

3.1.5 Single-Event Effects

Shielding has limited effectiveness as a mitigation for single-event effects, which can be induced by deeply penetrating high-energy particles (e.g., Galactic Cosmic Rays). The primary method for radiation hardness is to use parts insensitive to destructive SEE (e.g., single-event latchup, gate rupture, or burnout). For non-destructive effects (e.g., single-event upsets, transients, or functional interrupts), mitigation takes the form of fault isolation, detection, and recovery; error-detection and correction codes (EDACs); filtering circuitry; power-cycling; and other techniques to mitigate or eliminate any effect on broader system operation.

For single-event effects that have potential to propagate within a design, a Single Event Effect Criticality Analysis (SEECA) can be used to identify the single-event effect sensitivity and severity. Further information on event criticality is found at:

http://radhome.gsfc.nasa.gov/radhome/papers/seecai.htm. Alternative methods for propagating SEE effects in a design are open for development. Such considerations required in the analysis include time to detect failures, time to recover failures, system operations, and mission phases as in Radiation Single Event Effects (SEE) Impact on Complex Avionics Architecture Reliability, NASA/TM-2019-220269.

Single-Event Effect Criticality Classes:

• Error-Functional - function may be unaffected by SEE (possibly by error-correction scheme, mitigation, or redundancy); frequent events may be acceptable.

• Error-Vulnerable - function where low probability for SEE is required; response by mitigation or characterization and acceptance of risk is permissible.

• Error-Critical - function where SEE are unacceptable; part technology, design, or usage must eliminate any probability of error to a specified degree.

In the following sections, requirements for EEEE parts are categorized by single-event effect type. Many missions use linear energy transfer (LET) thresholds to establish compliance. In general, higher LET particles are rarer in a space environment and certain bounds are established to determine the types of testing or test data required. Supporting data will address the part's system application (bias voltage, frequency, etc.) to determine applicability to these requirements.

Notes on common LET bounds:

• Proton-induced secondaries in silicon and SiC materials have LET < 20 MeV-cm2/mg.

This range includes the vast majority of fluence for a typical mission.

• The natural space environment has a precipitous decline in particle populations that exhibit LET > 37 MeV-cm2/mg.

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• Ion strikes that follow angular paths through sensitive volumes may have a higher effective LET (37 MeV-cm2/mg could have an effective 75 MeV-cm2/mg).

SEE are largely dependent on component design and architecture, rather than lot-to-lot process variability. Data is generally applicable (“flight lot applicable”) where there have been no changes to the semiconductor manufacturer mask set, foundry, or process.

3.1.5.1 Non-Destructive SEE (e.g., SEUs, SETs, SEFIs, SHEs)

LSERD-182 All systems shall be designed to avoid or tolerate errors due to non-destructive SEEs.

LSERD-183 Non-destructive SEEs shall not cause permanent loss of function.

Note: There is no arbitrary LET threshold (LETth) imposed for devices that show non-destructive single event effects. However, devices with LETth >20 MeV-cm2/mg are recommended to reduce system risks.

3.1.5.2 Destructive SEE

LSERD-200 No single-event effect shall cause permanent failure in a critical component.

Note: Critical components are components that are Category 1SC or 1 in the Severity Categories Table 4.2-1 in the Landsat 10 (Landsat 10) Instrument Mission Assurance Requirements (IMAR), L10-SMA-REQ-0013. For reference, 1SC are failures that could cause a catastrophic event such as the loss of life, permanently disabling injury to personnel, or facility loss/destruction. Category 1 are failures (and failures with failure recoveries) that could result in mission loss or minimum mission success criteria not being achievable. Reentry capability failures are included in Cat 1. Single Point Failures are to be categorized as error-critical and must be documented.

LSERD-201 No single-event effect shall cause permanent degradation that adversely affects system performance or reliability.

3.1.5.2.1 LET Threshold Requirements for SEL

LSERD-496 All devices shall have LET thresholds for SEL greater than 37 MeV-cm2/mg.

Note: For devices with SEL thresholds > 75 MeV-cm2/mg, or those built from a technology not susceptible to SEL, no additional analysis is necessary.

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3.1.5.2.2 Single-Event Burnout and Gate Rupture (SEB, SEGR)

3.1.5.2.2.1 SEB and SEGR Requirements

All devices that are susceptible to single-event burnout (SEB) or single-event gate ruptures (SEGR) should be properly derated in the susceptibility and performance analysis.

LSERD-971 Single-event burnout and gate rupture shall not cause permanent damage to, or performance degradation of, the flight system.

3.1.5.2.3 LET Threshold Requirements for SEB and SEGR

LSERD-225 All parts that may be susceptible to SEB or SEGR shall have a SEGR and

SEB threshold LET > 37 MeV-cm2/mg when biased at 133% of the application Drain-to-Source Voltage (VDS), Collector-to-Emitter Voltage (VCE), or Reverse Voltage (VR).

LSERD-226 Equivalently, all parts that may be susceptible to SEB or SEGR shall be operated at no more than 75% of the highest VDS, VCE, or VR at which the part does not experience SEB or SEGR at an LET > 37 MeV-cm2/mg.

LSERD-227 All power MOSFETs shall have a SEGR threshold LET > 37 MeV-cm2/mg when biased at 167% of the application Gain-to-Source Voltage (VGS) (off) condition and at 133% of the application VDS.

LSERD-228 Equivalently, all power MOSFETs shall be operated at no more than 60% of the VGS (off) at which the part does not experience SEGR.

3.2 MICROMETEOROID AND ORBITAL DEBRIS (MMOD) ENVIRONMENTS

The Landsat 10 SOWs for the spacecraft and instrument address the NASA program requirements for the mission to meet NASA Procedural Requirements for Orbital Debris Mitigation, NPR 8715.6. The NPR addresses limiting orbital debris when in orbit and also addresses how to responsibly remove space assets from low-Earth orbit (LEO). This section addresses a different concern related to Micrometeoroid and Orbital Debris (MMOD).

This section addresses how the spaceborne asset will minimize risk to mission success from a science perspective. The risk to mission success is stand-alone from reliability mission success and other formal mission success calculations. This is Landsat 10’s effort to provide a higher probability of science return while operating in the proposed MMOD environment. Orbital Debris Mitigation, NPR 8715.6 solely addresses the risk of impacting other spaceborne assets and ground debris.

Appendix C provides further details and clarifications.

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LSERD-247 Spacecraft and instrument materials and components shall be designed to remain operable within their performance specifications over the mission design lifetime in the micrometeoroid and space debris environments in the operational orbit.

NOTE: This requirement applies to all observatory orientations, including operational and safehold

LSERD-250 The spacecraft and instruments shall include shielding against micrometeoroid and orbital debris that is compliant with Figure 3.2-1 or Table 3.2-1 for instruments and for non-de-orbit critical spacecraft components.

NOTE: Does not apply to the following items:

• Optical apertures

• Stand-alone thermal (radiator) apertures (Still applies to thermal sides of electronic boxes)

• Solar array

• Thruster apertures

• Mechanisms

• Redundant harnesses that are physically separated from each other by a circumferential shield; the redundant harnesses may be wrapped together in a larger bundle and still be considered separated

NOTE: Both wake and nadir sides of the observatory may use a standard 15-layer MLI (~0.0241g/cm2 areal density) and do not require the enhanced MMOD MLI; the remaining four sides require the enhanced MMOD MLI.

NOTE: Appendix C provides additional information and guidelines

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

Figure 3.2-1 Design Curves for Protection against Penetration of 1.4mm 14.5 km/s particles

Table 3.2-1 Blanket Areal Density (g/cm2) at Given Wall Thickness, 14.5 km/s

Separatio n

Distance

Blanket Areal Density (g/cm2) at a Given Wall Thickness (cm)

S (cm) 0.15 0.20 0.25 0.30 0.35 0.40 0.45 0.50 0.60 0.70 0.80

1.0 cm 0.073

0.062

0.053

0.046

0.040

0.029

0.019

0.011

1.5 cm 0.074

0.638 0.055

0.048

0.042

0.037

0.027

0.018

0.010

2.0 cm 0.083

0.068

0.059

0.052

0.045

0.040

0.035

0.026

0.017

0.010

2.5 cm 0.076

0.064

0.057

0.050

0.044

0.038

0.034

0.025

0.017

0.010

3.5 cm 0.086

0.070

0.061

0.054

0.048

0.042

0.037

0.032

0.024

0.016

0.009

5.0 cm 0.078

0.066

0.058

0.052

0.046

0.040

0.036

0.031

0.023

0.016

0.009

7.5 cm 0.073

0.063

0.055

0.050

0.044

0.039

0.035

0.030

0.023

0.015

0.009

10 cm 0.070

0.061

0.054

0.049

0.043

0.038

0.034

0.030

0.022

0.015

0.009

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3.3 OBSERVATORY MAGNETIC FIELDS

LSERD-396 The spacecraft and instruments shall not exhibit any malfunction, degradation of performance, or deviation from their performance requirements during exposure to natural and self-induced magnetic fields encountered between earth's surface and final mission orbit

3.4 ATOMIC OXYGEN

Initial analysis shows an accumulated Atomic Oxygen (AO) fluence of 7.462E19 atoms/cm2 for Landsat 10's 63-month mission. This value constitutes the upper bound of AO fluence for Landsat 10 orbit parameters assuming the 653 km orbit vs 705 km. The magnitude of AO fluence will erode about 0.09 mils of a regular Kapton blanket with the +2σ scenario and assume 100% in the ram direction. For solar array panel, it would have an AO fluence of 2.375E19 atoms/cm2 due to its rotation. Reference Figure 3.4‑1 below for the flux curve.

Figure 3.4-1 Atomic oxygen flux (atoms/cm2 /month) for 63-month (5- year mission plus

3-month commission with accumulated AO fluence for + 2σ solar activities

LSERD-399 The spacecraft and instruments shall meet performance requirements during exposure to atomic oxygen fluence from Figure 3.4-1 experienced over the mission lifetime in the environments in the operational orbit.

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3.5 OBSERVATORY CHARGING FROM ALL SOURCES

For high-inclination, polar low Earth orbits (polar LEO), like Landsat 10, Mitigating In-Space Charging Effects – A Guideline, NASA-HDBK-4002 is the guideline to be followed for mitigating in-space charging effects. Low Earth Orbit Spacecraft Charging Design Standard, NASA-HDBK-4005 and Low Earth Orbit Spacecraft Charging Design Handbook, NASA- HDBK-4006 provide guidance for the low inclination portions of polar LEO. However, those two documents apply to spacecraft at orbital inclinations that seldom encounter the auroral zones. Per Figures 1 and 2 (page 12) of NASA-HDBK-4002, Landsat 10 is a medium risk for surface charging hazards and low risk for internal charging hazards.

The requirements herein are based on Class S grounding guidance provided in Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads, and Flight Equipment, NASA-STD- 4003 and the surface charging and internal charging hazards specified in NASA-HDBK-4002 for the Landsat 10 orbit.

The following resistivity definitions apply to electrostatic requirements of this section:

Table 3.5-1 Resistivity Definitions

Term Definition Insulative and Dielectric Materials

Has a surface resistivity of at least 1 x 10^12 Ω/sq or a volume resistivity of at least 1 x 10^11 Ω-cm.

Dissipative Materials Have a surface resistivity equal to or greater than 1 x 10^5 Ω/sq but less than 1 x 10^12 Ω/sq or a volume resistivity equal to or greater than 1 x 10^4 Ω-cm but less than 1 x 10^11 Ω-cm.

Conductive Materials Have a surface resistivity less than 1 x 10^5 Ω/sq or a volume resistivity less than 1 x 10^4 Ω-cm.

Charging Potential Materials with a view to space and capable of obtaining differential potential of 100 volts or great or any internal material capable of obtaining a differential potential greater than 400 volts. The material must be nominally greater than 6 cm2 in surface area.

Grounding The requirements herein are based on Class S grounding guidance provided in NASA-STD-4003A. Application of the grounding schema will be design dependent.

The following requirements apply to all mission phases in the plasma environment (specified in Appendix A) at worst-case cold operating temperature during all phases of solar illumination.

LSERD-402 The spacecraft and instruments shall operate without performance degradation due to the surface charging and deep charging environments.

Note: This requirement covers any items not explicitly covered by the following requirements.

The providers are assumed to parse requirements to the proper hardware to meet the overall requirement. Without a known design, Landsat 10 cannot explicitly cover the spacecraft or instruments. This requirement will be verified by the completion of the lower-level requirements linked to it.

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3.5.1 Observatory Charging Potential

Differential potentials of 100 volts or greater pose a discharge risk in the high plasma densities encountered in LEO. Internal volumes that have low plasma densities pose a discharge risk for differential potentials above 400 volts. The maximum safe area criteria are a practical limit to allow small ungrounded areas without the need for a waiver.

Refer to:

Overview of Plasma Interactions Section, Low Earth Orbit Spacecraft Charging Design Handbook, NASA-HDBK-4006, page 14, Appendix A, A.1 Poisson Equation, paragraph a., for 100-volt discharge threshold in LEO environment.

Spacecraft charging (21 July 2008), ECSS-E-ST-20-06C, Section C.9.4.1. Metallic discharge:

grounded conductor, for 100-volt discharge threshold for inverted gradient charging.

Mitigating In-Space Charging Effects – A Guideline, NASA-HDBK-4002, page 66, paragraph 5.2.3.2.1, Grounding Conductive Elements for the area limit. Refer to Mitigating In-Space Charging Effects – A Guideline, NASA-HDBK-4002, page 44, Section 4.5.1, Dielectric Surface Breakdowns, paragraph a. for 400-volt discharge threshold.

LSERD-404 Observatory surfaces greater than 6 cm2 in area with a view to space shall maintain a surface potential of less than 100 volts with respect to the observatory primary structure, adjacent grounds or conductors, and adjacent dielectric surfaces.

Note: The 6 cm2 area limit is a compromise to allow ungrounded ties, tapes, fasteners, etc. of limited area that are not near sensitive electronics.

LSERD-999 Observatory surfaces greater than 6 cm2 in area without a view to space, including surfaces of internal volumes, shall maintain a surface potential of less than 400 Volts with respect to the observatory primary structure, adjacent grounds or conductors, and adjacent dielectric surfaces.

Note: The 6 cm2 area limit is a compromise to allow ungrounded ties, tapes, fasteners, etc. of limited area that are not near sensitive electronics.

3.5.2 Observatory Charging Electric Field

LSERD-412 All observatory surfaces and materials greater than 6 cm2 in area with a charging potential shall maintain an internal electric field of less than 2 X 107 V/m.

Rationale: Electric fields of approximately 2 X 107 V/m and greater pose a discharge risk. The maximum safe area criteria are a practical limit to allow small ungrounded areas without the need for a waiver.

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Note 1: The 6 cm2 area limit is a compromise to allow ungrounded ties, tapes, fasteners, etc. of limited area that are not near sensitive electronics.

Note 2: Sharp edges, points, and imperfections will enhance the electric field at a given potential compared to smooth surfaces and large radius transitions. Refer to Mitigating In-Space Charging Effects – A Guideline, NASA-HDBK-4002, page 32, section 4.1.5 Breakdown Voltage, for maximum allowable electric field.

LSERD-1162 All observatory surfaces and materials greater than 6 cm2 in area with a charging potential shall maintain an electric field of less than 1 X 107 V/m between a visible surface dielectric and an exposed grounded conductor.

3.5.3 Observatory Grounding to Chassis Ground Plane (CGP)

LSERD-417 Individual conductive layers of multi-layer insulation blankets, unused cable conductors and cable conductors isolated by relays, unconnected traces, metal integrated circuits (IC) lids, cans (IC, transistor, or crystal), and any metal in close proximity to or connected to sensitive electronics anywhere on the observatory shall be grounded back to the spacecraft Observatory Grounding to Chassis Ground Plane (CGP) through a resistance not to exceed (109/ (area in cm2)) ohms with the following exceptions.

a) Ungrounded metal with surface area < 6.5 cm2 is allowed if it is enclosed by > 60 mil equivalent aluminum shielding (net from 4 pi steradian ray trace) and is maintained at a temperature greater than -30 C at all times.

b) Anodize fasteners, hinges, pins, etc. do not provide a ground path due to the insulative surface.

c) Grounds should be provided for components that are equipped with explicit ground points or ground pins, such as ICs with ground pins for lids.

Note: Deviations and/ or waivers may be granted for the following cases:

• For areas protected by < 60 mils aluminum equivalent shielding where the modeling and analysis indicates low risk.

• For components with demonstrated geosynchronous earth orbit (GEO) flight heritage where the on-orbit performance indicates low risk, and sufficient technical information describing the materials and geometry of the ungrounded metal and the potential nearby discharge sources and/or targets is provided and also indicates low risk in the particular Landsat 10 observatory configuration.

LSERD-423 All components shall provide a ground path that bypasses any bearings.

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3.5.3.1 Grounding of Dielectric Films and Bulk Materials

LSERD-426 All dielectric films exceeding 6 cm2 in area shall have a grounded conductive coating on at least one side.

Rationale: The charging rate is governed by absorbed flux, material conductivity, and capacitance from the material to spacecraft ground. Dielectric films, such as sheet Kapton, that are not in intimate contact with a grounded surface may have very low capacitance to ground, and therefore charge much more rapidly than comparable grounded films, leading to high surface potentials and possible discharge.

Note: Dielectric films with a direct view to space (e.g., thermal blankets) must have the conductive coating on the exposed side of the outer layer if a non-conductive base material is used.

LSERD-429 All dielectric bulk materials with a charging potential and with bulk resistivity less than 1 x 109 ohm-cm shall be grounded with a Class S bond per Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads, and Flight Equipment, NASA-STD-4003 to the nearest accessible ground point.

LSERD-431 All dielectric bulk materials with a charging potential and with bulk resistivity greater than 1 x 109 ohm-cm shall have a conductive coating or wrap grounded to the nearest accessible ground point.

Note:

a) This requirement is specifically not applicable to PCBs that have a ground or power plane within 80 mils of both surfaces, are behind at least 60 mils aluminum equivalent shielding, and remain above -30 C regardless of power state.

b) PCBs must have at least one ground or power plane in all PCB areas

– open areas (with no ground or power plane) are not allowed.

3.5.3.2 Thermal Blankets

LSERD-435 Thermal blankets shall have ground connections as defined in Table 3.5.3.2-

1.

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Table 3.5.3.2-1 Table Ground Bond Connections per Blanket Area

Blanket Surface Area (cm2) Quantity of Ground Bond Connections - Minimum <6 0, use of conductive tape to establish an adjacent blanket as a ground point is preferred <200 Use of conductive tape to establish an adjacent blanket as a ground point is allowed, all VDA layers must be grounded

>200 and <10,000 2 >10,000 and < 30,000 3

Each Additional 20,000 1 additional

LSERD-453…