Attachment I - LNext Space Envrioments Requirements Document (LSERD).pdf

DOORS EXPORT

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

LNEXT-SYS-REQ-0009, Revision - Landsat Next, Code 426

Landsat Next Space Environments Requirements Document (LSERD) ecavanag CMO Release Stamp

LSERD LNEXT-SYS-REQ-0009

Revision -ii

Landsat Next Space Environments Requirements Document

(LSERD)

Signature Pages Prepared by:

Electronic Signature in TDMS

12/13/2022

Michael Pryzby Date Landsat Next Systems Engineer NASA/GSFC, Code 426

Approved by:

Mark Edison Date Landsat Next Deputy Observatory Manager

Wen-Ting Hsieh Date Landsat Next Payload Manager

Joy Henegar-leon Date Landsat Next Payload Technical Manager iii

Evan Webb Date Landsat Next Systems Manager NASA/GSFC, Code 599

James Pontius Date Landsat Next Project Manager NASA/Goddard, Code 426 iv

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

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 Next 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 Next Project Office, Code 426 Attention: Configuration Management Office Greenbelt, Maryland 20771 v

Change History Log Revision Effective Date Description of Changes

- 12/13/2022 LNEXT-CCR-0024– Initial Release vi

List of TBDs/TBRs Hyperlink to TBx Location Summary Ind.

Name/Org. Due Date

TBx-1 Section

1.2.4 Figure is TBD Michael

Pryzby SRR/MDR vii

Table of Contents SIGNATURE PAGES .................................................................................................................... II

CM FOREWORD ......................................................................................................................... IV

CHANGE HISTORY LOG ............................................................................................................ V

LIST OF TBDS/TBRS .................................................................................................................. VI

TABLE OF CONTENTS ............................................................................................................. VII

LIST OF FIGURES .................................................................................................................... VIII

LIST OF TABLES ...................................................................................................................... VIII

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 Landsat Next 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) ...9

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

viii

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 1.2.4-1 Landsat Next Requirements Tree Figure 2.4-1 Landsat Next Operations Concept Figure 3.2-1 MMOD Shielding Blanket Thickness 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 Fontheim Auroral Electron Flux Figure A-2 Mission Average Integral Omnidirectional Current Density Projected Through a Flat Surface, (nA/cm2) Figure A-3 Worst-Case Omnidirectional Integral Flux Projected Through a Flat Surface

(nA/cm2) Figure C-1 Overall Description of the Proposed Process Figure C-2 Omnidirectional Orbital Debris Flux Curve for the Landsat 9 Orbit Figure C-3 Directionality of the Orbital Debris Threat to Landsat Next Figure C-4 Design Curves for Protection against Penetration of 1.8mm 14.7 km/s particles

List of Tables Table 3.2-1 Blanket Areal Density (g/cm2) at Given Wall Thickness (cm) Table 3.5-1 Resistivity Definitions ix

Table 3.5.3.2-1 Table Ground Bond Connections per Blanket Area Table A-1 Nascap-2k Parameters for Surface Charging Table A-2 Fontheim Auroral Electron Flux Table A-3 Mission Average Integral Omnidirectional Current Density Projected Through a Flat Surface, nA/cm2 Table A-4 Mission Average Integral, Omnidirectional Number Flux (Electrons/cm2-sec )

Table A-5 Worst-Case Omnidirectional Integral Flux Projected Through a Flat Surface

(nA/cm2) Table A-6 Worst-Case Integral Omnidirectional Flux (Electrons/cm2/s) Table C-1 Directional Orbital Debris Flux (#/m2-yr) and Velocity (km/s) for the Landsat

Next Orbit Table C-2 Numerical Values for the Data Plotted in Figure C-4 Table C-3 Increase in Effective Wall Thickness from Layers of Kevlar

1.0 INTRODUCTION

1.1 PURPOSE

The Landsat Next Space Environments Requirements Document (LSERD) establishes a set of Level 2 requirements for safe operations of the Landsat Next Observatories (spacecraft and instruments) in their predicted orbit and space environment.

1.2 SCOPE

1.2.1 Identification

This Landsat Next (LNext) Space Environments Requirements Document (LSERD) sets forth the general space environment for the LNext 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 Next spacecraft and 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 one or more instruments. This term is used mainly in reference to testing in order to differentiate between the spacecraft bus without instruments and the integrated satellite with instruments.

The expression instrument provider refers to the organization / company delivering the instruments for integration on the spacecraft. The term is used interchangeably with the instrument contractor or sensor subcontractor with respect to Landsat Next. For the purposes of this document, the instrument provider is treated as a single entity, although, in reality, the integrating team will probably include representatives from many organizations.

The expression separately mounted instrument components refers to each part of an instrument that is separately mounted onto the spacecraft by the spacecraft contractor. Where an instrument is divided into multiple pieces but is mounted onto the spacecraft via a single baseplate, it is not considered ‘separately mounted’ instrument components in this document.

The term "(TBD)," which means "to be determined," applied to a missing requirement means that, depending on the applicability of the requirement, the instrument contractor or spacecraft contractor should propose the resolution of the specification, with the final determination to be made by the government.

The term "(TBR)," which means "to be reviewed," means that the requirement is subject to review for appropriateness by both contractors and the government, and subject to revision. The instrument and spacecraft contractors are responsible for compliance with the requirement as if the "TBR" notation did not exist. The "TBR" merely provides an indication that the value is more likely to change in a future modification than requirements not accompanied by a "TBR."

1.2.4 Document Overview

This LSERD is controlled and maintained by the Landsat Next 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.

The following Figure 1-1 shows the relationship between requirements documents for the Landsat Next mission.

TBD

Figure 1.2.4-1 Landsat Next Requirements Tree

• Environmental Requirements Document (ERD) – parent document of the radiation requirements specification – the ERD contains a subset of the requirements (shall statements) listed in this document.

• Radiation Environment Description Document (REDD) – child document of radiation requirements specifications - radiation environment modeling and predictions will be captured in a separate document to accommodate mission design / life changes.

• As-Design / as-built parts lists (ADPL/ABPL) – will capture Parts Control Board (PCB) assumptions and approvals for requirements verifications (with references to test results, analysis, dependencies, etc.).

• Mission Assurance Requirements (MAR) - A radiation hardness assurance program (RHAP) or ionizing radiation control plan (IRCP) that is generated in accordance with MAR, Mission Assurance Implementation Plan (MAIP), and EEEE Parts Control Plan (PCP).

1.2.5 Rationale and Notes

Throughout this document, there are paragraphs which are provided for information, rationale, or clarification, but do not represent requirements. These are shown in italics. 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.

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 applicable documents of the revision and release date shown. 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 NEXT 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 Section Reference NASA-STD-4003 Electrical Bonding for NASA

Launch Vehicles, Spacecraft, Payloads, and Flight Equipment

A w/CHANGE 1 Section 4.1.4, 4.1.6, 5

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

This document table was generated from the Landsat Next Referenced Documents List Draft Rev -

1.3.2 Landsat Next 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 Spacecraft and LandIS Statements of Work.

Document Number Title Revision Section Reference

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 Page 31 Section 4.1.5, Page 60 Section 4.4.1, paragraph 5.2.3.2.1, Page 64 Section 5.2.4.3, Page 71 Section 5.2.5.1

NASA-HDBK-4006 Overview of Plasma Interactions A Page 14 Appendix A

NASA/TM-2018-220074 "Guidelines for Verification Strategies to Minimize RISK Based on Mission Environment, - Application and -Lifetime (MEAL)"

N/A

NASA/TM-2019-220269 "Radiation Single Event Effects (SEE) Impact on Complex Avionics Architecture Reliability"

N/A

NEPP, 2009 “Proton Single Event Effects (SEE) Guideline,” NASA Electronic Parts and Packaging (NEPP)

N/A

2015 IEEE NSREC

Short Course

2015 IEEE Nuclear and Space Radiation Effects Conference Short Course Notebook

Document Number Title Revision Section Reference

2017 IEEE NSREC

Short Course

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

2017 IEEE NSREC

Short Course

"Total Ionizing and Non-Ionizing Dose Radiation Hardness Assurance"

ECSS-E-ST-20-06C Spacecraft Charging (15 May 2019) 1 Page 24: Section 3.5.1

IEEE Trans. Nucl. Sci "Statistical Model Selection for TID Hardness Assurance," Ladbury, R.,

et. al., IEEE Trans. Nucl. Sci., vol.

56, no. 6, pp. 3354-3360

N/A

IEEE Trans. Nucl. Sci “Damage Correlations in Semiconductors Exposed to Gamma, Electron, and Proton Radiation," Summers, G.P., et. al., IEEE Trans.

Nuc. Sci., vol. 40, no. 6, pp. 1372- 1379, 1993.

N/A

IEEE Trans. Nucl. Sci "Inclusion of Radiation Environment Variability in Total Dose Hardness Assurance Methodology," Xapsos, M.A., et. al., IEEE Trans. Nucl. Sci., vol. 64, no. 1, pp. 325-331, Jan.

2017.

N/A

IEEE Trans. Nucl. Sci "SEE Characterization of Vertical DMOSFETs:

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

Nucl. Sci., vol. 50, no.6, pp. 2341- 2351, 2003

N/A

IEEE Trans. Nucl. Sci Recommended Test Conditions for SEB Evaluation of Planar Power DMOSFETs," Liu, S., et al., IEEE Trans. Nucl. Sci., vol.

55, no.6, pp. 3122-3129

N/A

MIL-STD-883 Microcircuits Test Standards L Page 28: Section 3.5.5

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

This document table was generated from the Landsat Next Referenced Documents List Draft Rev -

2.0 MISSION OVERVIEW

2.1 MISSION STATEMENT

Landsat Next, consistent with United States (U.S.) law and government policy, will continue the Landsat program's acquisition, archival, and distribution of multi-spectral imagery affording global, synoptic, and repetitive coverage of the Earth's land surfaces 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 National Aeronautics and Space Administration (NASA) recognized the need to assemble a team of experts from within both agencies to evaluate how to 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 was 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 – Release 2.0.” These investigations 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 mass held in the National Satellite Land Remote Sensing Data Archive (NSLRSDA) at 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.

The goal of Landsat Next is to continue the acquisition, archival, and distribution of multi-spectral imagery affording global, synoptic, and repetitive coverage of the Earth's land surfaces 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 Next continues the long-running partnership of NASA and USGS, with NASA providing the space and launch segments and USGS providing the ground system, providing the longest continuous global record of the Earth’s surface. 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. The Landsat data archive constitutes the longest continuous moderate-resolution record of the global land surface as viewed from space.

2.3 MISSION OBJECTIVES

Landsat Next has these major mission objectives:

• Collect and archive moderate resolution multispectral image data and thermal image data, affording seasonal coverage of the global landmass for a continuous period of not less than five (5) years.

• Ensure that Landsat Next 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 to permit studies of land cover and land use change over multi-decadal periods.

• Ensure Landsat Next is responsive to critical emerging user needs and applications as characterized by periodic assessment, currently the User Needs for the Sustainable Land Imaging Program July 2018, Release 2.0, and identified by the operational requirements for collection, processing, archiving, and distribution of land surface data to the United States Government and other users.

• Distribute Landsat Next data products to the general public on a nondiscriminatory basis.

2.4 MISSION IMPLEMENTATION

NASA and USGS each have specific responsibilities for Landsat Next and will deliver the major elements to the overall mission. NASA will provide the Space and Launch Segments of Landsat Next, and USGS will provide the Ground System and mission operations. NASA/GSFC will provide overall Landsat Next project management, mission system engineering, and mission assurance during development and will transition the mission to USGS following on-orbit commissioning.

The Landsat Next post-launch nominal mission operations concept is shown graphically in the Figure below.

Figure 2.4-1 Landsat Next Operations Concept

2.4.1 Space Segment

NASA/GSFC will provide the Space Segment via competitive procurements for the science instruments and the spacecraft bus. The Space Segment will consist of a constellation of three observatories flying in coordinated sun-synchronous orbits at 653km altitude, each with nominally identical spacecraft and instrument suites. The observatories will be equally spaced in the orbit, providing in aggregate a six-day ground repeat period at the equator.

The instrument suites will each provide 26 spectral bands with a maximum ground sampling distance of 10m-60m dependent on the spectral band, covering a swath on the ground of approximately 164km on a ground track defined by a world-wide reference system known as

WRS-3.

2.4.2 Ground Segment

The Landsat Ground Segment currently supports mission operations for Landsats 8 and 9.

Landsat Next takes advantage of developments on these missions and will upgrade and expand the current systems to accommodate Landsat Next. The Landsat Ground Segment consists of the Ground System (GS) 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) and NASA ACCESS Space relay (i.e. TDRSS) and government/commercial ground station, NASA/GSFC Conjunction Assessment and Risk Analysis (CARA) and its NASA/GSFC Flight Dynamics Facility, along with other external interfaces.

The MOC provides the primary means to control and monitor the Landsat Next constellation.

The Landsat Next Flight Operations Team (FOT) at the MOC performs mission planning and scheduling, command and control, health and status monitoring, orbit and attitude maintenance, performance analysis, onboard memory management, and flight and ground software maintenance. The FOT utilizes MOC functionality to detect, investigate and resolve spacecraft anomalies and monitor the instrument image collections from the onboard constellation and generate special image collections. The MOC ingests, processes and archives data via the GN.

The GN includes geographically dispersed ground station resources for mission execution, and includes both the Landsat Ground Network (LGN) and a wideband or cloud-based data routing capability to transfer both mission data to DPAS, and TT&C data to the MOC. The LGN consists of US Government-owned, international, and commercial ground stations, and provides communication capability for each observatory of the constellation for commanding and housekeeping data via S-Band two-way links. The LGN will also receive mission data from each observatory via high rate Ka-band downlinks.

The DPAS ingests, processes, and archives LNext mission data from the GN. 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 Next science products, via a public-facing web portal for receiving data products. The DPAS is located at USGS Earth Resources Observation and Science (EROS) near Sioux Falls, South Dakota.

The USGS will lead overall Landsat Next 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 perform planning, scheduling, and observatory operations activities during Landsat Next mission readiness testing.

2.4.3 Launch Segment

The Launch Segment will provide the assets and services associated with the Launch Vehicle (LV) and the constellation-to-LV integration. This will include the LV; all Launch Vehicle- Ground Support Equipment (LV-GSE), property, and facilities to integrate the constellation to the LV and verify their integration; and prelaunch testing with ground-based functions. The launch vehicle and launch site are TBD. The three observatories comprising the Landsat Next constellation will be launched together on the same launch vehicle.

Sections 2.0-2.4.3 were generated using Landsat Next Common Boilerplate (LNEXT-MGMT- DESC-0005) Rev B

3.0 ENVIRONMENTAL CONDITIONS

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

3.1 CHARGED PARTICLE RADIATION ENVIRONMENT

The LNext 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 EEEE parts shall function after exposure to the 95th percentile mission total ionizing dose as specified in the LNext Radiation Environment Definition Document, LNEXT-SYS-DESC-0015, Figure 5.2.1-2.

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 specified in Radiation Environment Definition Document Figures 5.3.1-1 and 5.3.1-2 for the LNext mission for the appropriate material (most commonly silicon or gallium arsenide).

3.1.3 TID/DDD Testing

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

Flight lot testing is mandatory for commercial (non-DLA audited) parts, even where generic (non-lot-specific) data exist.

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.

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 LNext Radiation Environment Definition Document, 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 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.

• 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.

http://radhome.gsfc.nasa.gov/radhome/papers/seecai.htm

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 SEUs.

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.

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.

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 VDS, VCE, or 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 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 LNext SOWs for the observatory and instrument(s) address the NASA program requirements for the mission to meet NPR 8715.6B, NASA Procedural Requirements for Limiting Orbital Debris and Evaluating the Meteoroid and Orbital Debris Environments. 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 relative 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’s effort to provide a higher probability of science return while operating in the proposed MMOD environment. The NPR referenced above solely addresses the risk of impacting other spaceborne assets and ground debris.

Appendix C provides further details and clarifications.

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 attitudes, 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.

Notes:

1. This requirement 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

2. 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

3. Appendix C provides additional information and guidelines

Figure 3.2-1 MMOD Shielding Blanket Thickness

Table 3.2-1 Blanket Areal Density (g/cm2) at Given Wall Thickness (cm)

Separation

Distance

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

S (cm) 0.10 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 0.126 0.104 0.086 0.078 0.068 0.050 0.036 0.026

1.5 cm 0.118 0.104 0.090 0.078 0.070 0.062 0.048 0.034 0.022

2.0 cm 0.110 0.092 0.082 0.072 0.064 0.058 0.044 0.034 0.022

2.5 cm 0.120 0.104 0.090 0.078 0.070 0.062 0.056 0.042 0.030 0.022

3.5 cm 0.110 0.096 0.084 0.074 0.068 0.060 0.054 0.040 0.030 0.022

5.0 cm 0.120 0.100 0.090 0.082 0.072 0.064 0.058 0.050 0.040 0.030 0.020

7.5 cm 0.110 0.096 0.086 0.076 0.070 0.062 0.056 0.050 0.040 0.028 0.020 10 cm 0.124 0.106 0.092 0.084 0.076 0.068 0.062 0.054 0.048 0.040 0.028 0.020

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 LNext’s 63-month mission. This value constitutes the upper bound of AO fluence for LNext orbit parameters. 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.

3.5 OBSERVATORY CHARGING FROM ALL SOURCES

For high-inclination, polar low Earth orbits (polar LEO), like Landsat Next, NASA-HDBK- 4002B is the guideline to be followed for mitigating in-space charging effects. NASA-HDBK- 4005 and NASA-HDBK-4006A 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-4002B, LNext 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 NASA-STD- 4003A and the surface charging and internal charging hazards specified in NASA-HDBK-4002B for the LNext 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, LNext cannot explicitly cover the spacecraft or instruments. This requirement will be verified by the completion of the lower-level requirements linked to it.

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:

NASA-HDBK-4006A, page 14, Appendix A, “Overview of Plasma Interactions,” A.1 “Poisson Equation”, paragraph a., for 100-volt discharge threshold in LEO environment ECSS-E-ST-20-06C, Section C.9.4.1. "Metallic discharge: grounded conductor" for 100-volt discharge threshold for inverted gradient charging NASA-HDBK-4002B, page 66, paragraph 5.2.3.2.1, “Grounding Conductive Elements” for the area limit. Refer to NASA-HDBK-4002B, page 44, Section 4.4.1,"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, and they will maintain an electric field of less than 1 X 107 V/m between a visible surface dielectric and an exposed grounded conductor.

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.

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 NASA-HDBK-4002B, page 32, section 4.1.5 “Breakdown Voltage,” for maximum allowable electric field.

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.

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 Next observatory configuration.

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

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 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.

Table 3.5.3.2-1 Table Ground Bond Connections per Blanket Area

Blanket Surface Area (cm2) Quantity of Ground Bond Connections - Minimum

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

Each Additional 20,000 1 additional

LSERD-453 The linear distance from any point on the blanket as measured along the surface of the blanket to the nearest ground bond connection tab shall be equal to or less than 1 meter.

LSERD-454 The resistance along the length of any grounding strap as measured near the blanket connection point vapor deposited aluminum (VDA) on an inner blanket layer to the terminal end of the grounding strap on chassis/structure shall be less than or equal to 10 ohms.

LSERD-455 After installation, the resistance along the length of any grounding strap as measured near the blanket connection point VDA on an inner blanket layer to structure ground shall be less than or equal to 10 ohms.

LSERD-456 Blanket surface resistivity shall be less than 5 x 106 ohms/square.

Note: Kapton 275 XC or Kapton 160XC is suggested as the optimum base material for the outer blanket layer for exposed blanketed surfaces.

Ref. NASA-HDBK-4002B, page 78, section 5.2.5.1, Thermal Blankets

3.5.4 Solar Array Surface Charging

LSERD-459 Solar arrays shall maintain surface voltages below 100 volts with respect to the observatory primary structure; adjacent grounds or conductors; and adjacent dielectric surfaces, maintain an internal electric field of less than 2 X 107 V/m in dielectric materials, and maintain an electric field of less than 1 X 107 V/m between a visible surface dielectric and an exposed grounded conductor.

Note: The junctions between cover glass to uninsulated electrical interconnects may present high electric fields if best practices are not followed. See NASA-HDBK-4002B and NASA-HDBK- 4006A.

If the requirements of LSERD-404 and LSERD-412 are not met for the solar array, a deviation may be granted based on:

1) Evidence of successful flight heritage in a very similar orbit. The heritage design must be identical at the panel level.

2) Test results from the original flight heritage design, to include clear evidence that sustained arcing cannot occur.

3) Analysis of the heritage and current designs indicating acceptable performance in the expected flight environment.

This information must demonstrate immunity to sustained arcing and no significant performance degradation, damage, or loss of function to the solar array due to solar cell damage or cover glass contamination from repeated arcing, or other induced effects from trigger arcs. In addition, it must be demonstrated that the observatory will not be adversely affected by trigger arc induced noise conducted or radiated into observatory systems, or contamination projected onto any observatory surface.

LSERD-466 Solar arrays shall be designed to prevent sustained arcing induced by electrostatic discharge (ESD) events, micrometeoroid impacts, or other plasma producing events on the both the front and back sides of the arrays.

Note: Current that can support sustained arcing can be sourced from the illuminated solar array or back fed from the solar array regulators.

Ref. NASA-HDBK-4002B, page 71, section 5.2.4.3, Solar Array Design Guidelines to Protect Against Space Charging and ESDs Ref. J. Michael Bodeau. "Test Approach to Assess Solar Array Diode Board Susceptibility to Sustained Arcing," Journal of Spacecraft and Rockets, Vol. 50, No. 6 (2013), pp. 1277-1287.

doi: 10.2514/1.A32442 Ref. J. Michael Bodeau, “Sustained Arc Qualification Test of an Array Diode Board,” 2012 Space Power Workshop, April 16-19, 2012.

3.5.5 Sensor Shielding

LSERD-485 Sensor cable shielding shall be extended to cover the sensor.

Note: The cable shield Faraday cage can be extended by attaching copper or aluminum tape directly to the cable shield and extending it to cover the sensor. If the sensor (temp sensor, proximity sensor, position sensor, for example) cannot be enclosed, the shield or tape should be extended far as possible to reduce the length of unshielded cable.

LSERD-488 Human Body Model Class 1A parts or better shall be used for electronics box external interfaces.

Note: HBM Class 0 parts may be damaged by small transients generated by ESD in cable dielectrics. The Human Body Model is defined in MIL-STD-883-3, Method 3015.9 Electrostatic Discharge Sensitivity Classification.

Appendix A Plasma Environment Specification

SPACE CHARGING ENVIRONMENT

This section presents the worst-case charged particle (electrons and protons) environments likely to be encountered from launch through operation in polar orbit. This section provides the environments only. These environments will be used to establish design and verification of space charging mitigation requirements for the subsystems, elements, and platform.

The two main categories of space charging environments, “surface charging” and “bulk charging” (also known as “internal” or “deep dielectric” charging), are represented here.

The term “integrated flux density” represents the total number of particles impinging on a flat plate per time per area that have energy greater than or equal to the energy value shown.

SURFACE CHARGING ENVIRONMENT MODEL PARAMETERS

Table A-1 includes plasma environment parameters for input to the NASCAP-2K surface charging code. NASCAP-2K predicts surface voltages and electric fields (among other outputs) in surface materials and can be used to verify compliance with relevant space charging requirements for surface materials.