LNEXT-SC-REQ-0011 Rev B Landsat Next SC SRD 20250124.pdf

DOORS EXPORT

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

LNEXT-SC-REQ-0011, Revision B Landsat Next, Code 426

Landsat Next Spacecraft Requirements Document (SRD)

LNext Reviewed – Not Subject to Export Control

LNext SRD LNEXT-SC-REQ-0011 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 Next Spacecraft Requirements Document (SRD) Signature/Approval Page

Prepared by:

Electronic Signature in TDMS Approval Date in TDMS Bryan Mattfeld Date Landsat Next Spacecraft Systems Engineer NASA/GSFC, Code 560

Approved by:

Cagatay Aymergen Date Acting Landsat Next Project Manager / Landsat Next Deputy Project Manager

NASA/GSFC, Code 426

Cagatay Aymergen Date Acting Landsat Next Observatory Manager NASA/GSFC, Code 426

Evan Webb Date Landsat Next Mission System Manager NASA/GSFC, Code 599 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 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 iv Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Change History Log

Revision Effective Date Description of Changes (Reference the CCR/SCoRe Approval Date)

- 05/01/2024 LNEXT-CCR-0084 – Initial Release A 11/12/2024 LNEXT-CCR-0091 – Update for draft SC RFP B 01/23/2025 LNEXT-CCR-0112 – Updated to include the latest versions of the Landsat Next Common Boilerplate, LNEXT-MGMT- DESC-0005 Rev D and Landsat Next Common Mission Architecture Diagram, LNEXT-SYS-DESC-0022 Rev A.

Updated two common definitions, "component" and "unit."

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

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

Name/Org.

Due Date

TBx-1 Section 3.1.3

LN_SRD-28 The observatory shall complete commissioning activities within 90 days (TBR) of launch.Rationale: The 90-day (TBR) requirement for observatory commissioning is meant to constrain the amount of time necessary before commencing operations even though handover to operations occurs 100 days after launch. It is desired that the spacecraft will finish the majority of activation within the first 30 days or sooner to allow the focus of commissioning effort to shift to the instrument activation and calibration. Note that the timeline for achieving final mission orbit will likely not be fully determined until the final year prior to launch, and the spacecraft commissioning may occur entirely at the parking orbit or at intermediate ascent altitudes.

In addition to the 90 days for commissioning, it is expected that 10 days will be necessary for handover.

MSE 08/24

TBx-2 Section 3.13.8

LN_SRD-661 Upon receipt of a valid cold start ground command, spacecraft operational flight software shall complete its startup within TBD seconds.

SC SE 08/24

TBx-3 Section 4.4

LN_SRD-1094 Each of the three (3) LNext Observatories shall be capable of initiating the mission power up sequence based upon receipt of a wakeup signal from the LV via the LV umbilical harness prior to separation if applicable to the Mission Payload design.

Note: The format of the wakeup signal is (TBD).

SC SE 08/24

TBx-5 Section 4.4

LN_SRD-1095 Each of the three (3) LNext Observatories shall be capable of providing a telemetry stream to the LV for interleaving with the LV telemetry stream during pre-launch and ascent via the LV umbilical harness if applicable to the Mission Payload design.

Note: The format of the observatory telemetry stream is (TBD). The LV will accommodate three separate telemetry streams generated by the Mission Payload at a rate per observatory not to exceed 2.0 kilobits per second.

SC SE 08/24

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

TBx-5 Section 4.5

LN_SRD-1100 The Mission Payload design, to include the three LNext Observatories, shall allow for periods with the Mission Payload Electrical Ground Support Equipment (EGSE) being disconnected from the LV umbilical interface following the completion of final Mission Payload closeout activities for up to 24 (TBR) hours during combined operations activities with the LV.

Note: This applies to the umbilical interfaces identified in Requirement LV-IDD-40. The number and duration of outage periods will be determined during the launch integration process.

SC SE 08/24

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Table of Contents SIGNATURE/APPROVAL PAGE ................................................................................................ II

CM FOREWORD ..........................................................................................................................III

CHANGE HISTORY LOG .......................................................................................................... IV

LIST OF TBDS/TBRS .................................................................................................................... V

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

LIST OF FIGURES ........................................................................................................................ X

LIST OF TABLES .......................................................................................................................... X

1.0 INTRODUCTION

1.1 Purpose

1.2 Scope

1.3 Related Documentation

1.3.1 Applicable Documents

1.3.2 Reference Documents

1.3.3 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 DESIGN LIFE

3.1 Mission Phases

3.1.1 Pre-Launch Phase

3.1.2 Launch and Early Orbit Phase

3.1.3 Spacecraft Commissioning Phase

3.1.4 Spacecraft Decommissioning Phase

3.2 Orbits

3.2.1 Injection Orbit

3.2.2 Operational Orbit

3.2.3 Disposal Orbit

3.3 Launch Vehicle Interface

3.4 Imaging Support

3.5 Spacecraft Operations

3.5.1 Spacecraft Modes

3.5.2 Operational Contingency Response

3.6 Redundancy and Single-Fault Tolerance Design Requirements

3.7 Fault Management

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

3.7.1 Telemetry Monitoring

3.8 Availability

3.9 Data Loss

3.10 Structural and Mechanical Systems

3.10.1 Mechanisms and Deployable Design Requirements

3.11 Thermal System

3.12 Electrical System

3.12.1 Energy Collection and Storage

3.12.2 Power Quality

3.12.3 Power Distribution and Monitoring

3.12.3.1 Power Distribution for Hazardous Functions

3.12.3.2 Power Fault Protection

3.12.4 Grounding

3.12.4.1 Power and Signal Returns

3.12.4.2 Class R Grounding

3.12.4.3 Electrostatic (Class S) Grounding

3.12.5 Flight Plugs

3.12.6 Harness and Connector Design

3.12.6.1 Signal and Power Routing

3.12.6.2 Connector Types and Identification

3.12.6.3 External Interfaces

3.13 Flight Software

3.13.1 Startup Read Only Memory (SUROM)

3.13.2 Deterministic Processing

3.13.3 Memory Operations

3.13.4 Memory Margin

3.13.5 Flight Software Patch Upload

3.13.6 Version Identifiers

3.13.7 Event Logging

3.13.8 Initialization

3.13.9 Commands

3.13.10 Stored Commands

3.13.11 State-of-Health (SOH) Telemetry

3.13.12 Ancillary Data

3.14 Firmware

3.14.1 Version Identifiers

3.15 Command and Data Handling System

3.15.1 General

3.15.2 Observatory ID

3.15.3 Commands

3.15.4 Mass Storage

3.15.4.1 General

3.15.4.2 Mission Data Files (MDFs)

3.15.4.3 Volume Table of Contents (VTOC)

3.15.4.4 Erase and Record Operations

3.15.4.5 Playback Operations

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

3.15.4.6 Mass Storage Data Integrity

3.15.4.7 Diagnostic and Test

3.16 Spacecraft Operational Autonomy

3.16.1 Activities

3.16.1.1 Activity Definition

3.16.2 Communications Activity Schedule (CAS)

3.16.2.1 Communications Activity Schedule Inputs

3.16.2.2 Communications Activity Schedule (CAS) Generation

3.16.2.3 Communications Activity Schedule Outputs

3.16.3 Imaging Plan

3.16.3.1 Long Term Acquisition Plan (LTAP) Collection Request (LCR) ..54

3.16.3.2 Special Collection Request (SCR)

3.16.3.3 Imaging Plan Generation

3.16.3.4 Imaging Plan Outputs

3.16.4 Ground-Directed Activity Timeline (GDAL)

3.16.4.1 GDAL Inputs

3.16.4.2 GDAL Generation

3.16.4.3 GDAL Outputs

3.16.5 Integrated Onboard Timeline (IOT)

3.16.5.1 IOT Generation

3.16.5.2 IOT Outputs

3.17 Attitude and Orbit Control

3.17.1 Maneuvers

3.17.1.1 Orbit Maneuvers

3.17.1.2 Attitude Maneuvers

3.17.2 Ephemeris

3.17.3 Spacecraft State Knowledge

3.17.4 Spacecraft Attitude Knowledge

3.17.5 Spacecraft Attitude Control

3.18 Propulsion System

3.19 Radio Frequency Telecommunications System

3.19.1 General

3.19.2 Narrowband (S-Band)

3.19.3 Wideband (Ka-Band)

APPENDIX A MISSION PAYLOAD REQUIREMENTS

A.1 Mechanical Interfaces A.2 Mission Payload Final Closeouts Prior to Encapsulation A.3 Mission Payload Propellant Loading and Contingency Offloading A.4 Airborne Interfaces A.5 Ground Interfaces

APPENDIX B ABBREVIATIONS AND ACRONYMS

APPENDIX C DEFINITIONS

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List of Figures Figure 2.4-1 Triplet Constellation Observatory Spacing On-Orbit Figure 2.4-2 Landsat Next Mission Architecture

List of Tables

Table 2.4.1-1 Mission Parameters Table 3.2.1-1 Injection Orbit Table 3.2.2-1 LNext Operational Orbit

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1.0 INTRODUCTION

1.1 PURPOSE

The Landsat Next Spacecraft Requirements Document (SRD) defines the requirements for the Landsat Next Spacecraft.

1.2 SCOPE

The Landsat Next Spacecraft Requirements Document (SRD) establishes a set of Level 3 requirements for the Landsat Next Spacecraft in support of the Landsat Next Project. As a Level 3 document, it contains functional and performance requirements.

The requirements in main body of the SRD provide requirements to be levied on each individual spacecraft such that each spacecraft must meet requirements as a standalone unit. The requirements in Appendix A outline functional and performance requirements applicable to the Mission Payload (including dispenser system).

The following definitions apply to the action verbs used in this requirements document:

• Shall: Directive – Compliance by the Contractor is mandatory. Any deviations from these contractually imposed mandatory requirements require the approval of the contracting officer.

• May: Permissive – Allows an option for compliance at the discretion of the Contractor or Government.

• Will: Informative – Designates the intent of the Government. Unless required by other contract provisions, noncompliance with the will requirements does not require approval of the contracting officer and does not require documented technical substantiation.

• Should: Advisory – Gives guidance to the Contractor for the performance of this contract.

• Must: Advisory – Notifies the Contractor of important relevant conditions that may be outside the context of this contract.

1.3 RELATED DOCUMENTATION

1.3.1 Applicable Documents

Document Number Title Revision Section Reference

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

- All

NASA-STD-4003 Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads, and Flight Equipment

A w/CHANGE 1 All

NASA-STD-5017 Design and Development Requirements for Mechanisms

B All

NASA-STD-8719.14 Process for Limiting Orbital Debris C All NASA-STD-8719.24 NASA Expendable Launch Vehicle

Payload Safety Requirements A All

NPR 8715.6 NASA Procedural Requirements for Orbital Debris Mitigation

E All

CCSDS 232.1-B-2 Communications Operation Procedure-1 Recommended Standard, Issue 2, All

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 Section Reference

Technical corrigendum 1

CCSDS 301.0-B-4 Time Code Formats Recommended Standard, Issue 4, Editorial Change

All

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

This document table was generated from the Landsat Next Common Reference and Applicable Documents List (LNEXT-MGMT-DESC-0008) Rev E.

1.3.2 Reference Documents

Document Number Title Revision GSFC-STD-1000 Rules for the Design, Development, Verification, and

Operation of Flight Systems H

Public Law #105-303 Commercial Space Act of 1998, Public Law #105-303, 112 Stat., 2843., H.R. 1702, 105th Congress, Adopted October 28, 1998

N/A

Public Law #102-555 Land Remote Sensing Policy Act of 1992, Public Law #102- 555, 106 Stat. 4163., H.R. 6133, 102nd Congress, Adopted October 28, 1992

N/A

CCSDS 133.1-B-3 Encapsulation Packet Protocol Recommended Standard, Issue 3, Editorial Change 1

NIMA TR8350.2 Department of Defense World Geodetic System 1984 Third Edition, Amendment 1

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

This document table was generated from the Landsat Next Common Reference and Applicable Documents List (LNEXT-MGMT-DESC-0008) Rev E.

1.3.3 Project Reference Documents

Document Number Title LNEXT-LS-ICD-0009 Landsat Next Mission Payload to Launch Services Interface Description

Document (LNext-LS IDD) LNEXT-SC-ICD-0002 Landsat Next Spacecraft Interface Requirements Document (SCIRD) LNEXT-SC-ICD-0007 Landsat Next Spacecraft to Ground Interface Requirements Document (SC-

GND IRD)

LNEXT-SC-PLAN-0055 Landsat Next Spacecraft Contract Data Requirements List (CDRL) and Data

Item Descriptions (DID) LNEXT-SMA-REQ-0015 Landsat Next Spacecraft Mission Assurance Requirements (SCMAR) LNEXT-SYS-DESC-0007 Landsat Next Geometric Error Budget (GEB) Support Presentation LNEXT-SYS-DESC-0016 Landsat Next Geometric Error Budget (GEB) LNEXT-SYS-DESC-0018 Landsat Next World Reference System-3 (WRS-3) LNEXT-SYS-PLAN-0027 Landsat Next Project Protection Plan (PPP) LNEXT-SYS-TN-0001 Systems Technical Memo - Debris Casualty Area (DCA) Suballocation for

Landsat Next LSDS-2389 Landsat Next Design Reference Case – 18 Days (DRC-18) Definition In this document, citations are assumed to be the latest version unless otherwise noted.

This document table was generated from the Landsat Next Common Reference and Applicable Documents List (LNEXT-MGMT-DESC-0008) Rev E.

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 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 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 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 July 2018, 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 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.

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 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 Next 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 Next 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 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; 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 Next 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 Next data products to the public freely on a nondiscriminatory basis.

2.4 MISSION IMPLEMENTATION

The Landsat Next mission is a constellation of three identical Observatories, referred to as “triplets” with equal orbit spacing. Each observatory images the full swath required to achieve global coverage, with the use of three observatories used to reduce revisit time for any scene.

The Figure below illustrates the 120 degree separation of the Observatories.

Figure 2.4-1 Triplet Constellation Observatory Spacing On-Orbit

This selected architecture of three observatories has been endorsed by the Landsat user communities and improves all temporal, spectral, and spatial capabilities by approximately twofold, while maintaining the historical Landsat data continuity. With all three observatories operating, imaging revisit will be every six (6) days, at the equator.

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Each agency has specific responsibilities and will deliver major elements to the overall mission.

NASA will provide the Space and Launch Segments of Landsat Next and the USGS will provide the Ground System and Operations. NASA/Goddard Space Flight Center (GSFC) will provide overall Landsat Next project management, mission system engineering and mission assurance during development and will transition the mission to USGS for science operations following on-orbit commissioning. Each agency in addition to specific responsibilities, will deliver the three major segments to the overall mission.

The three segments are the Space Segment (SS), the Ground Segment (GS), and the Launch Services Segment (LSS). They are described in more detail in the following sections and shown in the Landsat Next Operations Concept below.

Figure 2.4-2 Landsat Next Mission Architecture

(Generated from Landsat Next Common Mission Architecture Diagram LNEXT-SYS-DESC-0022 Rev A)

2.4.1 Space Segment

NASA/GSFC will provide the Space Segment via competitive procurements for the science instrument suite and the spacecraft bus. The Space Segment will consist of a constellation of three observatories flying in sun-synchronous, polar orbits at 653km altitude, each with nominally identical spacecraft and instrument suites. The three observatories will be launched on a single launch vehicle with a single dispenser provided by the spacecraft bus provider. The observatories will be transitioned to their final orbit alignments during commissioning. In aggregate, they provide a six-day ground repeat period at the equator. The spacecraft and instrument will be Class C. The Mission Class B is accomplished by providing redundancy at

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the Observatory level. The overall mission parameters needed to achieve global coverage are shown below in the Mission Parameters Table.

Table 2.4.1-1 Mission Parameters

A single Observatory will consist of the spacecraft and imaging sensors, referred to as the Landsat Next Instrument Suite (LandIS). Each observatory operates independently, and unique coordination is conducted by the ground as the observatories do not communicate with each other directly.

The LandIS instrument suite is a whiskbroom scanner which provides 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. The instrument suite will be capable of acquiring raw detector data for all spectral bands within 15 second to minimize the illumination changes within the observations. Data will be generated at a rate no greater than 1.8 Gbps. The LandIS provides internal calibration sources such as blackbody, a space port, stim lamps and solar diffuser. LandIS instrument suite provides compression prior to packaging the data for transfer to the spacecraft storage.

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. Space Segment simulator types include spacecraft, instrument suite, spacecraft and the LandIS instrument suite interface, launch vehicle interface and a combination of simulators integrated to be an Observatory. These simulators will have varying fidelity and capabilities to meet the mission needs.

2.4.2 Ground Segment

The Landsat Ground Segment 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

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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 Next MOC functionality will be used for command and control, health and status monitoring, orbit and attitude maintenance, performance analysis, onboard memory management, and flight and ground software maintenance. The MOC will support onboard autonomy for mission planning and scheduling. The MOC ingests, processes and archives data delivered via the GN. The Landsat Next Flight Operations Team (FOT) located at the MOC performs these functions and is responsible for monitoring the Observatories, 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 for each observatory of the constellation for commanding and housekeeping data via the S-Band and will receive science mission data from each observatory over a Ka-band downlink. The LGN is used to transfer commands and receive telemetry from each Observatory within the constellation and deliver it to the MOC.

The DPAS ingests, processes, and archives science data returned from each satellite via 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 data products, via a public-facing web portal for receiving data products.

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 support Landsat Next 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 Mission Payload-to-LV integration. The Launch Segment will include the LV; all Launch Vehicle-Ground Support Equipment (LV-GSE), property, and facilities to integrate the Mission Payload to the LV and verify its integration; and perform prelaunch testing with ground-based functions. The launch vehicle and launch site have not been specified and will not be until closer to Mission Phase C. The three observatories comprising the Landsat Next Mission Payload, including the Observatory dispenser, will be launched together on the same launch vehicle.

The candidate launch vehicles for the launch of the Landsat Next Mission Payload are the Falcon 9 Full Thrust, the United Launch Alliance Vulcan Centaur and the Blue Origin New Glenn. This list will be updated in the future to reflect additions and deletions to the candidate launch vehicles presently available under the NASA Launch Services II Contract.

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

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3.0 DESIGN LIFE

LN_SRD-8 The spacecraft (SC) shall be designed to meet all operational requirements for the 5.25-year design mission life.

Rationale: In support of the extensive database of global land use it is paramount that Landsat Next (LNext) provide a meaningful mission life to support this effort. The Landsat Next Spacecraft needs to provide operational support for at least five years to meet its Level 1 and 2 requirements and objectives. LNext Mission life is 5 years. This requires a high reliability bus design to meet the overall constellation of a minimum of two observatories operating for five years in a degraded 3:2 constellation, providing 9 day revisit at minimum. Mission life begins after the observatory commissioning period. Each spacecraft must satisfy the mission design requirement itself, independent of the other spacecraft within the constellation.

3.1 MISSION PHASES

3.1.1 Pre-Launch Phase

LN_SRD-13 The spacecraft shall be integrated with Government-furnished Landsat

Instrument Suite (LandIS) in compliance with the Landsat Next Spacecraft Interface Requirements Document (SCIRD), LNEXT-SC-ICD-0002.

LN_SRD-17 The spacecraft shall be designed to perform aliveness tests of the spacecraft and integrated LandIS while mated to the LV.

Rationale: To identify any possible ground handling problems and provide confidence in a successful mate to the LV, the spacecraft needs to perform some aliveness tests. This includes supporting instruments and other subsystems.

LN_SRD-969 The spacecraft shall monitor battery voltage continuously up to launch vehicle fairing closeout and via the umbilical through launch.

LN_SRD-19 The spacecraft shall be capable of test firing the thrusters with dry gas at the assembled spacecraft level.

Rationale: This supports thruster phase and functional tests during Integration and Test (I&T).

3.1.2 Launch and Early Orbit Phase

LN_SRD-23 The total LNext mission payload mass, fully fueled in launch configuration, including all 3 observatories and dispenser system, shall not exceed 8000 kilograms (kg).

Rationale: The launch configuration includes the observatories loaded with propellant and any ballast required for meeting LV maximum center of gravity (CG) offset requirements. The mass margins specified in Goddard Open-Learning Design (GOLD) Rules for Design, Development, Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Verification and Operation of Flight Systems, GSFC-STD-1000, Gold Rule 1.06 will be managed at the mission level. Instrument NTE mass is defined in the SCIRD per SCIRD-160.

3.1.3 Spacecraft Commissioning Phase

LN_SRD-28 The observatory shall complete commissioning activities within 90 days

(TBR) of launch.

Rationale: The 90-day (TBR) requirement for observatory commissioning is meant to constrain the amount of time necessary before commencing operations even though handover to operations occurs 100 days after launch. It is desired that the spacecraft will finish the majority of activation within the first 30 days or sooner to allow the focus of commissioning effort to shift to the instrument activation and calibration. Note that the timeline for achieving final mission orbit will likely not be fully determined until the final year prior to launch, and the spacecraft commissioning may occur entirely at the parking orbit or at intermediate ascent altitudes. In addition to the 90 days for commissioning, it is expected that 10 days will be necessary for handover.

3.1.4 Spacecraft Decommissioning Phase

LN_SRD-32 The spacecraft shall protect the near-Earth space environment by complying with NASA Procedural Requirement for Orbital Debris Mitigation, NPR 8715.6.

Rationale: This requirement minimizes potential for future orbital debris, minimizes non-operational lifetime on orbit, minimizes Earth impact debris footprint, and has a system capable of performing any required maneuvers at end of mission design life. Inputs will be provided in support of the orbital debris assessment report (ODAR) and end of mission plan (EOMP).

LN_SRD-33 The spacecraft shall be designed for controlled re-entry in compliance with

Process for Limiting Orbital Debris, NASA-STD 8719.14.

Rationale: This is in lieu of a natural decay and uncontrolled re-entry.

LN_SRD-34 The reliability of postmission disposal maneuver operations for controlled re-entry shall be 0.9 or greater at the end of mission design life.

Rationale: This requirement is derived from Process for Limiting Orbital Debris, NASA-STD- 8719.14C, requirement 4.6-4 (a). To ensure that a controlled re-entry can be performed successfully at the end of the mission; the systems necessary to complete this maneuver must have a high reliability. Calculations to determine controlled re-entry probability are governed by the Landsat Next Spacecraft Contract Requirements Document List (CDRL) and Data Item Descriptions (DID), LNEXT-SC-PLAN-0055 SA-12 document.

LN_SRD-846 Each spacecraft shall have a maximum debris casualty area (DCA) of 55 m^2.

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Rationale: The DCA is defined per the Systems Technical Memo - Debris Casualty Area (DCA) Suballocation for Landsat Next, LNEXT-SYS-TN-0001 .

LN_SRD-42 The spacecraft shall maintain attitude control for all propulsive maneuvers including the last disposal maneuver.

Rationale: Attitude control is required for all maneuvers. The de-orbit maneuvers may be long and are likely to be affected by atmospheric drag. Attitude control is necessary to minimize impact dispersions.

3.2 ORBITS

3.2.1 Injection Orbit

LN_SRD-37 The spacecrafts shall begin commissioning after being injected into the orbit defined in Table 3.2.1-1

Rationale: The LV will target direct insertion of the LNext spacecraft into the operational orbit with the allowable LV dispersions.

Table 3.2.1-1 Injection Orbit

Parameter Value Tolerance Mean Semi-major Axis 7026.33 km

(IO-1-1)

± 20 km

(IO-1-2)

Mean Apogee Radius 7045

(IO-2-1)

± 20 km

(IO-2-2)

Inclination 97.9835 °

(IO-3-1)

± 0.1 deg

(IO-3-2)

Argument of Perigee 90.0 °

(IO-4-1)

N/A

(IO-4-2)

Mean Local Time (MLT) of Descending Node 10:10

(IO-5-1)

± 1.0 minute

(IO-5-2)

This table was made from the Landsat Next Common Operational Orbit Tables LNEXT-SYS-DESC-0036 Draft Rev -

3.2.2 Operational Orbit

LN_SRD-39 The spacecraft shall operate in a Sun-synchronous, near circular, frozen orbit with the orbital parameters as defined in the LNext Operational Orbit Table 3.2.2-1.

Rationale: The orbit defined in this requirement produces an orbit that follows the Worldwide Reference System-3 (WRS-3) Grid as defined in the Landsat Next World Reference System-3

(WRS-3), LNEXT-SYS-DESC-0018.

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Table 3.2.2-1 LNext Operational Orbit Parameter Value Allowable Variation Equatorial Altitude 653 km

(OO-1-1)

± 1.0 km

(OO-1-2)

Inclination (True – of – Date) 97.9835 °

(OO-2-1)

± 0.1°

(OO-2-2)

Mean Eccentricity (Brouwer-Lyddane) 0.00116

(OO-3-1)

<=0.00125

(OO-3-2)

Across track, ground track error at the descending node

0.0 km

(OO-4-1)

± 5.0 km

(OO-4-2)

Mean Local Time at Descending Node 10:10

(OO-5-1)

± 5 minutes

(OO-5-2)

Ground Track Spacing at Equator 151 km

(OO-6-1)

Mean Semimajor Axis 7026.33 km

(OO-7-1)

± 1.0 km

(OO-7-2)

Mean Argument of Perigee (deg) 90.0°

(OO-8-1)

Altitude Range 651 to 681 km

(OO-9-1)

This table was made from the Landsat Next Common Operational Orbit Tables LNEXT-SYS-DESC-0019 Rev -

3.2.3 Disposal Orbit

LN_SRD-41 The spacecraft shall achieve a perigee altitude of 50-km or less at the end of the final re-entry burn.

Rationale: To minimize the dispersions associated with the re-entry footprint, it is necessary to achieve a re-entry angle of attack large enough to ensure capture and prevent any atmospheric skipping of the observatory.

3.3 LAUNCH VEHICLE INTERFACE

LN_SRD-1072 Each observatory shall provide limited, essential telemetry to the launch vehicle via a single umbilical connection while in the launch configuration at a not to exceed (NTE) rate of 2 kbps per observatory.

Rationale: Each individual observatory will provide telemetry at 2kbps or less through an individual connector to the umbilical to enable monitoring of critical telemetry prior to separation. A total of 3 connectors will be needed for the mission payload to support the 3 spacecraft telemetry streams.

3.4 IMAGING SUPPORT

LN_SRD-49 The spacecraft shall continuously provide all imaging and calibration activities required to meet mission objectives as defined in the Landsat Next Design Reference Case – 18 Days (DRC-18) Definition, LSDS-2389.

LN_SRD-50 The spacecraft shall support acquisition of off-nadir images, up to 15 degrees cross-track.

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Rationale: For rapidly changing conditions on the earth, LNext will be able to image up to 15 degrees off-nadir left or right of the orbital track.

LN_SRD-51 The spacecraft shall support the acquisition of at least 9 minutes of continuous off-nadir imaging within a single orbit period.

Rationale: The distance across the conterminous United States is approximately 18 scenes.

Continuous long-duration off-nadir imaging will likely not be used extensively during operations.

LN_SRD-52 The spacecraft shall support the acquisition of at least 36 minutes of continuous nadir-pointing daylight imaging within a single orbit period.

LN_SRD-53 The spacecraft shall support the acquisition of at least 15 minutes of continuous nighttime imaging within a single orbit period.

Rationale: Note: that there will always be a gap of no imaging sequences between daylight only imaging sequences or between daylight and nighttime imaging sequences, where the imaging needs to be able to start without a need to restabilize the thermal conditions prior to an imaging sequence to start.

3.5 SPACECRAFT OPERATIONS

3.5.1 Spacecraft Modes

LN_SRD-56 The spacecraft shall be capable of being commanded into any operational mode or sub-mode.

Rationale: Allows for operational flexibility. Also clarifies valid ground commandable transitions. (e.g., Not possible to command the spacecraft into the launch mode). The full suite of available modes and sub-modes will be defined by the spacecraft vendor.

LN_SRD-1224 The spacecraft shall have a power-positive, thermally safe, control mode

(safe hold mode) with the following characteristics:

(1) Be capable of staying in this state indefinitely without ground command.

(2) Employ the minimum hardware set required to maintain a safe attitude.

Rationale: Safe hold mode is defined as part of Goddard Open-Learning Design (GOLD) Rules for Design, Development, Verification and Operation of Flight Systems, GSFC-STD-1000, Gold Rule 1.17. The spacecraft must be able to maintain safe hold mode for prolonged periods of time (>1 year) such that it may span multiple seasonal transitions/environments.

LN_SRD-57 The spacecraft shall autonomously transition from launch mode to safe hold mode upon separation from the Launch Vehicle (LV)

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Rationale: The spacecraft needs to be able to achieve a safe operating condition at separation from the LV.

LN_SRD-60 The spacecraft shall autonomously transition from any mode to the safe hold mode upon detection of faults that threaten the health and safety of the spacecraft.

Rationale: Upon the occurrence of a major fault the spacecraft must transition to a configuration that safes the spacecraft and instruments and slews to an attitude that ensures observatory power margins and thermal safety for an indefinite period of time without intervention from the ground.

Safe Hold Mode is defined as part of Goddard Open-Learning Design (GOLD) Rules for Design, Development, Verification and Operation of Flight Systems, GSFC-STD-1000, Gold Rule 1.17.

LN_SRD-61 The spacecraft shall transition out of the safe hold mode only via ground command.

Rationale: To ensure the safety of the observatory, it is critical that the spacecraft does not exit safe hold mode autonomously. Safe Hold Mode is defined as part of Goddard Open-Learning Design (GOLD) Rules for Design, Development, Verification and Operation of Flight Systems, GSFC-STD-1000, Gold Rule 1.17.

3.5.2 Operational Contingency Response

LN_SRD-63 The spacecraft shall be capable of indefinitely maintaining its health and safety without ground support.

Rationale: In the event of a failure in established communications, whether onboard the spacecraft or in the ground system, the spacecraft can respond appropriately without ground intervention. This may be achieved through autonomous fault management sequences, which result in the spacecraft initiating safing sequence(s), or other self-management activities which preserve observatory health and safety such as autonomous transition to Sun-pointing, or load-shed activities to power off non-essential units in the event of a fault. This state must be maintainable through the entirety of the 5.25 year life of the mission.

LN_SRD-65 The spacecraft shall continuously monitor its state-of-health (SOH) data, detect out-of-limit conditions, and respond to these out-of-limit conditions by configuring an appropriate safe state for the instrument and/or spacecraft systems.

LN_SRD-66 The spacecraft shall have the capability to detect, flag, and report out-of-limit conditions or other abnormal affects caused by external sources of Radio- Frequency Interference (RFI), particularly in the Radio Frequency (RF) communications, Global Positioning System (GPS), attitude control, and power subsystems, within 10 minutes of occurrence.

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Rationale: NASA's proactive response to space asset protection is that it is desirable for spacecraft to report unusual events that might indicate malicious as well as accidental/inadvertent activity. As an example, if an SC receiver experiences a very high input signal around the mission frequency, this may indicate an attempt to interrupt normal operations. By capturing the time of the event, it is possible to identify the area of the occurrence. LNEXT-SYS-PLAN-0027 Landsat Next Project Protection Plan will define relevant threats and mitigations that the mission will address. It is desired to detect and react quickly to any interference events and recover within an orbit whenever feasible.

3.6 REDUNDANCY AND SINGLE-FAULT TOLERANCE DESIGN REQUIREMENTS

LN_SRD-68 The spacecraft shall be designed such that no single credible failure of any mechanisms, thermal, and power systems that support the mission data chain permanently preclude the spacecraft from meeting all mission data acquisition, processing, storage, and wideband transmission requirements throughout the design life.

Rationale: Credible is defined as having a probability of occurrence greater than 0.001 for on-orbit technical risks and 0.000001 for safety risks at end of lifetime per Table 4.2.1-2 Likelihood Rankings in the Landsat Next Spacecraft Mission Assurance Requirements (SC MAR), LNEXT-

SMA-REQ-0015.

LN_SRD-1144 The spacecraft shall be designed such that there are no single credible failures within any system required to execute controlled re-entry.

LN_SRD-878 The mass storage system shall include 50% additional memory to be used as redundant, cold spare memory in the event of memory failures.

Rationale: This redundant memory is additional memory which is necessary to meet all storage, lifetime, and capacity including accounting for end of life sizing (EOL) requirements as specified in Section 3.15.4.

LN_SRD-69 The spacecraft shall preclude the propagation of spacecraft component, spacecraft subsystem, or LandIS failures, to other spacecraft components, spacecraft subsystems, or to LandIS.

Rationale: This includes precluding failure propagation to redundant sides of a system.

LN_SRD-74 The spacecraft shall be capable of switching between the primary and the backup unit for redundant systems by command.

Rationale: This should allow the spacecraft to switch back and forth between all primary and backup units through command.

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LN_SRD-75 The spacecraft shall explicitly identify for redundant systems, which unit is operational in SOH telemetry.

Rationale: The configuration of each subsystem should be clearly delineated in telemetry and should not be inferred based on current draw.

LN_SRD-76 Except for cross-strapped redundant power and signal services, the spacecraft shall route redundant electrical harness wires in a physically separated harness bundle.

Rationale: To protect from any credible event, which may cause the loss of a functional path, and to reduce the likelihood of the loss of the alternate or redundant path.

3.7 FAULT MANAGEMENT

LN_SRD-80 In the event of a failure that meets established failure criteria, the spacecraft shall autonomously switch to the backup component of redundant systems only as required to ensure observatory safety, and to enable the observatory to achieve safe hold mode.

Rationale: To ensure survivability in either the safe hold mode or the Earth-point sub mode, the spacecraft should be able to switch to a back-up element of a redundant system. There is no requirement for autonomous switching to support nominal imaging operations.

LN_SRD-81 The spacecraft shall preclude autonomously switching to a known failed redundant component.

Rationale: Switching to a known failed component is likely to be a threat to health and safety.

LN_SRD-83 The spacecraft shall be capable of over-riding autonomous functions via ground command. (Note: This does not include one-time protective hardware devices such as electrical fuses, etc.)

Rationale: The ground may need to override or suspend an autonomous response to ensure observatory safety.

LN_SRD-84 The spacecraft shall report autonomous reconfigurations in SOH and stored state-of-health (SSOH) telemetry.

Rationale: Whenever the state of a hardware configuration is changed, it needs to be tracked and reported in telemetry.

LN_SRD-85 The spacecraft shall send a command to alert LandIS to an unsafe condition upon detection of any specified off-nominal spacecraft conditions.

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Rationale: The LandIS must be alerted of any spacecraft mode changes from nominal to off-nominal or detected power, attitude, field of view fault, etc., problems so the instrument can elect to perform a self-protection as necessary.

LN_SRD-1131 The spacecraft shall place the LandIS in a known configuration in the event of any failures in the execution of operational autonomy.

Rationale: This does not preclude operational autonomy functions from coordinating with fault management (FM) to achieve the known and safe configuration for the instrument.

LN_SRD-218 The spacecraft shall be capable of activating a stored command sequence to perform fault management functions when errors are detected (i.e., a defined threshold is exceeded).

LN_SRD-970 The spacecraft shall override and/or interrupt any operational autonomy functions whenever errors are detected if necessary to perform fault management functions.

LN_SRD-1132 The spacecraft shall perform onboard autonomy functions on a non-interference basis with the fault management functions.

Rationale: It is critical that fault management functions always take precedence over operational autonomy and no fault management functions are either modified or acted upon by operational autonomy functions to ensure the health and safety of the observatory.

LN_SRD-219 The spacecraft shall maintain fault management function thresholds in tables in the flight software.

LN_SRD-227 The spacecraft shall perform table-driven fault management functions to check thresholds and persistent values on specific data items.

Rationale: Driving to define a table-driven approach for managing fault management.

LN_SRD-1061 The spacecraft shall disable fault management functions via critical command.

Rationale: Allow for certain fault management functions to be disabled during non-thermal vacuum (TVAC) I&T activities to facilitate testing.

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3.7.1 Telemetry Monitoring

LN_SRD-88 The spacecraft shall autonomously respond to spacecraft and LandIS SOH telemetry out-of-limit violations by execution of pre-defined, ground-configurable stored command sequences that safe the component, spacecraft, LandIS, or observatory.

Rationale: To provide a response based on changing states within the spacecraft.

LN_SRD-89 The spacecraft shall monitor pre-defined components, instrument(s), and subsystems, and support a ground-configurable telemetry monitoring capability, which can trigger a stored command sequence.

Rationale: The monitoring of SOH parameters in telemetry, such as power, temperature, operations of sub-systems, provides the information necessary to permit the spacecraft to take autonomous actions. Telemetry Monitors (TMONs) are useful in operations and troubleshooting to look at specific telemetry (TLM) points and execute actions when they exceed a specific value. This telemetry monitoring (TMONs) will be changed by ground command during operations. Either the SOH parameter or the action taken as a result of a change in the parameter will be changed by command. They are programmable from the ground to provide flexibility in addressing nominal and anomalous operations.

LN_SRD-92 The spacecraft shall allow for definition and modification of telemetry monitors on orbit via table upload.

Rationale: As telemetry monitors (TMONs) are programmable from the ground, they provide additional operations flexibility and support during troubleshooting.

LN_SRD-93 The spacecraft telemetry monitor…