LNEXT-SC-ICD-0002 Rev B Landsat Next Spacecraft Interface Requirements Document (SCIRD) 20250130.pdf

DOORS EXPORT

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

LNEXT-SC-ICD-0002, Revision B

Landsat Next, Code 426

Landsat Next Spacecraft Interface Requirements Document (SCIRD)

LNext Review – Not Subject to Export Control

1/30/2025

LNEXT SCIRD LNEXT-SC-ICD-0002 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 Interface Requirements Document

(SCIRD)

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 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 3.2.5.1.1

SCIRD-709 The spacecraft shall receive housekeeping data packets from LandIS at a rate of TBD kbps while the LandIS is in nominal configuration.

SC SE LandIS

PDR

TBx-2 Section 3.2.5.1.2

SCIRD-742 The high-speed data bus shall utilize a GbE channel, consisting of three (TBR) GbE links to transfer data from the LandIS to the Spacecraft.

SC SE LandIS

PDR

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

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

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

TABLE OF CONTENTS .............................................................................................................. VI LIST OF FIGURES ..................................................................................................................... VII LIST OF TABLES ....................................................................................................................... VII

1.0 INTRODUCTION

1.1 Purpose

1.2 Scope

1.3 Related Documents

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 INTERFACE REQUIREMENTS

3.1 General

3.2 SUBSYSTEM REQUIREMENTS

3.2.1 Mechanical Interface

3.2.1.1 Reference Frames

3.2.1.2 Volume and Dimension

3.2.1.3 LandIS Field of View

3.2.1.4 Mass Properties

3.2.1.5 Stowed and Critical Clearances

3.2.1.6 LandIS Access, Integration & Deintegration

3.2.1.7 LandIS Mounting

3.2.1.8 LandIS Dynamics

3.2.1.9 Instrument Mounted Spacecraft Components (IMSC) Sensor Accommodation

3.2.2 Orbit Control, Pointing and Alignment

3.2.2.1 Pointing Stability and Jitter

3.2.2.2 Alignment

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3.2.2.3 Spacecraft State Knowledge

3.2.2.4 Disturbance Torques

3.2.3 Thermal Interface Requirements

3.2.3.1 Operating and Survival Heaters and Temperatures

3.2.3.2 Mechanical Mounting Interface Temperature Monitoring

3.2.3.3 Thermal Isolation

3.2.4 Electrical Interface Requirements

3.2.4.1 Electrical Interfaces

3.2.4.2 Power Allocation

3.2.4.3 Main Bus Power

3.2.4.4 Power Distribution Interfaces

3.2.4.5 Spacecraft Class R Grounding Interface

3.2.4.6 Electrical Harness and Connectors

3.2.5 Command and Data Handling

3.2.5.1 Data Bus(es)

3.2.5.2 Time-of-day Pulse

3.2.5.3 Discrete Commands

3.2.5.4 Discrete Telemetry

3.2.6 Contamination Control

APPENDIX A ABBREVIATIONS AND ACRONYMS

APPENDIX B DEFINITIONS

List of Figures

Figure 2.4-1 Triplet Constellation Observatory Spacing On-Orbit Figure 2.4-2 Landsat Next Mission Architecture Figure 3.2.1.1.2-1 Observatory Orbital Reference Frame Figure 3.2.1.3-1 LandIS Fields of View, YZ Plane Figure 3.2.1.3-2 LandIS Fields of View, XZ Plane Figure 3.2.1.3-3 LandIS Fields of View, XY Plane Figure 3.2.1.8-1 Mass Acceleration Curve Figure 3.2.1.9-1 IMSC Footprint Figure 3.2.2.4-1 Torque Limit Figure 3.2.2.4-2 Deployment Torque Limits Figure 3.2.5.1-1 LandIS Functional Data Buses Figure 3.2.5.1-2 LandIS nominal instrument data block structure

List of Tables

Table 2.4.1-1 Mission Parameters Table 3.2.1.8-1 Component Design Loads (Break Points)

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1.0 INTRODUCTION

1.1 PURPOSE

The Landsat Next Spacecraft Interface Requirements Document (SCIRD) sets forth the initial interface requirements imposed between the spacecraft and the Landsat Next Instrument Suite (LandIS), with a focus on the spacecraft side of the interfaces, across the many subsystems, from Mechanical to Command and Data Handling.

1.2 SCOPE

Throughout this document references will be made to the LandIS and to the spacecraft.

References to the LandIS refer to the LandIS and apply only to the LandIS. References to the "Spacecraft" or "SC" applies only to the Landsat Next spacecraft.

The requirements that are indicated with "Reserved" text are applicable to the interface;

however, they are not allocated directly to that side of the interface.

The Landsat Next Interface Requirements documents (SCIRD, LIRD) are controlled and maintained by the Landsat Next Project office. The SCIRD is levied on the spacecraft contractor.

The LIRD is levied on the LandIS contractor. These IRDs are maintained at Level 3 and will contain sibling tracing to tie the SC interface requirements and LandIS interface requirements.

1.3 RELATED DOCUMENTS

1.3.1 Applicable Documents

Document Number Title Revision Section Reference EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

- TBD

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 NASA-STD-4003 Electrical Bonding for NASA Launch

Vehicles, Spacecraft, Payloads, and Flight Equipment

A w/CHANGE 1

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

CCSDS 301.0-B-2-S Time Code Formats Blue Book, Issue 2 CCSDS 701.0-B-3-S Advanced Orbiting Systems, Networks and

Data Links: Architectural Specification Blue Book, Issue 3

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

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

1.3.3 Project Reference Documents

Document Number Title LNEXT-SYS-DESC-0016 Landsat Next Geometric Error Budget (GEB) 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

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.

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

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

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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 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

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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 INTERFACE REQUIREMENTS

3.1 GENERAL

The Landsat Next Interface Requirements are maintained across two documents; namely the SCIRD and LIRD. They are controlled and maintained by the Landsat Next Project office. The SCIRD is levied on the spacecraft contractor. The LIRD is levied on the LandIS contractor.

These IRDs are maintained at Level 3 and will contain sibling tracing to link the SC interface requirements and LandIS interface requirements. Changes to a requirement with a sibling trace will require review/update of the corresponding sibling requirement and approval of both parties of the interface.

Since the Project is early in the mission, and the LandIS procurement will precede the Spacecraft procurement, the strategy is to have the LandIS contractor provide an Instrument Interface Description (IDD) Document. The IDD will be used to better inform and update these interface requirements.

For references to specific definitions and terms, refer to Appendix B.

3.2 SUBSYSTEM REQUIREMENTS

3.2.1 Mechanical Interface

3.2.1.1 Reference Frames

This SCIRD addresses multiple reference frames including the mechanical body, orbital reference, and optical reference frames. All reference frames use a right-hand, orthogonal, body fixed, XYZ coordinate system.

The mechanical body frame is a reference frame which is primarily used during assembly, integration, and test (AI&T). The reference frame is typically defined with the +X axis in the vertical direction as the mission payload would sit atop a launch vehicle. The +Y and +Z are defined to specific payload attributes. Where possible, +Z will be in the Nadir direction. The +Y /+Z plane is typically defined at a physical interface between items.

The orbital reference frame is primarily used for operations and flight dynamics. The orbital reference frame is typically defined about a space vehicle’s center of mass. +X is usually in the velocity vector. +Z is typically in the nadir direction. And +Y completes the right-hand rule.

The optical reference frame is used in the jitter, pointing, and alignment analysis, plus some ACS/GNC/FD analysis. +Z is the primary optical boresight of the Instrument. The XY plane is assumed to be the detector face. If they align, the +X direction would be in the velocity vector and +Y would complete the right-hand rule.

For Landsat Next, the two reference systems (mechanical and orbital) will have different origins and have axes in different directions. Attitude Control System (ACS), Guidance and Navigation, and Control (GNC), Flight Dynamics (FD), and Operations (Ops) will all be responsible for translating and rotating the coordinate systems for their analysis. The center of mass will be defined in the mechanical body frame reference.

3.2.1.1.1 Mechanical Body Frames

The mechanical body frames are a right-hand, orthogonal, body fixed, XYZ coordinate system.

The mechanical body frames are generally used during integration and test (I&T).

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The spacecraft contractor will include the translation between the SC mechanical body frame origin and the LandIS mechanical body frame origin on the Mechanical Interface Drawing.

3.2.1.1.1.1 Observatory Mechanical Body Frame

The dispenser mechanical Interface Control Document (ICD) will list the three Observatory mechanical body reference frames consistent with the three deployment systems.

The Observatory Body Frame is a right-hand, orthogonal, body-fixed XYZ system with +Y / +Z plane at the SC to Dispenser deployment interface plane, and the +Z direction pointing to orbit nadir, the +X direction in the LV velocity vector (or up when sitting on the dispenser adapter / vertical config), and the +Y completing the right-handed, orthogonal coordinate system. Given the orientation of the observatories on the dispenser, the Y and Z axes will not all match the Dispenser or Mission Payload Body frames without some rotation and/or translation.

3.2.1.1.1.2 Spacecraft Mechanical Body Frame

The Spacecraft body frame will be coincident with the observatory mechanical body reference frame.

The Spacecraft Body Frame is a right-hand, orthogonal, body-fixed XYZ system with +Y / +Z plane at the SC to Dispenser deployment interface plane, and the +Z direction pointing to orbit nadir, the +X direction in LV velocity vector (or up when sitting on the dispenser adapter / vertical config), and the +Y completing the right-handed, orthogonal coordinate system. Given the orientation of the observatories on the dispenser, the +Y and +Z axes will not all match the Dispenser or Mission Payload Body frames without some rotation and/or translation.

3.2.1.1.1.3 LandIS Mechanical Body Frame

LandIS Body Frame is a right-hand, orthogonal, body-fixed XYZ system with +Y / +Z plane at the SC to LandIS interface plane, and the +Z direction pointing in the orbit nadir direction, the +X direction in the LV velocity vector (or up when sitting on the dispenser adapter / vertical config), and the +Y completing the right-handed, orthogonal coordinate system.

Given mechanical body frames reference physical attributes, the origin of this frame is usually the centerline of a bolt hole on a fixed kinematic bracket and then along the theoretical interface plane between the LandIS bracket and the SC mounting.

3.2.1.1.2 Orbital Reference Frame

The Observatory Orbital Reference Frame is a right-handed, orthogonal, XYZ coordinate system defined such that the origin is at the Observatory center of mass (CM), the +Z axis points toward Geocentric Nadir, the +X axis is coplanar with both the +Z axis and the Observatory inertial velocity vector (and is in the general direction of the spacecraft inertial velocity vector), and the +Y axis completes the right-handed, orthogonal coordinate system, as shown in Figure 3.2.1.1.2- 1.

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Figure 3.2.1.1.2-1 Observatory Orbital Reference Frame

3.2.1.2 Volume and Dimension

SCIRD-139 The spacecraft shall accommodate a LandIS dynamic envelope of 1.95m

(76.77 in) in the Z axis by 1.37m (53.94 in) in the Y axis by 2.0 m (78.74 in) in the X axis in the launch configuration.

Note: The axes are per the LandIS mechanical body reference frame.

Rationale: The LandIS dimensional envelope constraints include the LandIS deck, IMSC and extends from the length of the flexures to the SC mounting point. A separate on-orbit volume requirement will need to be defined in the IRD if LandIS incorporates a deployable device into its operational configuration. The deployed envelope will not obstruct any observatory fields of view. Axes referenced are for the observatory body frame defined in Section 3.2.1.1.1.1

3.2.1.3 LandIS Field of View

SCIRD-153 The spacecraft shall provide the LandIS unobstructed optical and thermal

Fields of View (FOVs) as shown in Figure 3.2.1.3-1, Figure 3.2.1.3-2, and Figure 3.2.1.3-3.

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Figure 3.2.1.3-1 LandIS Fields of View, YZ Plane

Figure 3.2.1.3-2 LandIS Fields of View, XZ Plane

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Figure 3.2.1.3-3 LandIS Fields of View, XY Plane

3.2.1.4 Mass Properties

The not-to-exceed mass defined by the LandIS includes the fully assembled instrument, LandIS-unique flexures, instrument-unique fastening hardware, thermal blankets, connectors (instrument-side only), LandIS deck, and intra-instrument harness (with connectors), as well as any other items that constitute any part of a deliverable instrument, not including the IMSC hardware.

SCIRD-160 The spacecraft shall accommodate a total not-to-exceed 625 kg mass, including the LandIS and all associated hardware, intra-LandIS harnessing, any attached external harnessing extending to the spacecraft connector interface bracket, IMSC, and project-controlled instrument reserve mass.

Rationale: The total NTE mass of 625 kg given in this requirement that must be accommodated by the spacecraft includes the instrument mass, 25 kg of project-controlled instrument reserve mass and the IMSC mass defined in SRD-208. Note that IMSC mass remains a spacecraft budgeted item.

Total Mission Reserve allocations are managed by the Government and are not to be considered part of the allocated mass.

SCIRD-169 The spacecraft shall accommodate LandIS (including IMSC) having a stowed center of mass offset from the center point of the LandIS to Spacecraft interface plane (YZ plan) -662.94 mm in the X-direction, 19.812 mm in the Y-direction, 69.85 mm in the Z-direction.

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Note: The moments and products of inertia are to be measured or calculated using coordinates based on the spacecraft axes but passing through the LandIS center of mass.

SCIRD-172 Spacecraft moments and products of inertia values shall be accurate to within

± 5% for calculated values and the lesser of 1.5% or 300 kg-cm² for measured values.

3.2.1.5 Stowed and Critical Clearances

SCIRD-179 The spacecraft shall accommodate the LandIS so as to maintain adequate clearance between the LandIS and surrounding structures in order to provide access to LandIS mounting hardware, access to LandIS connectors, and space for LandIS interfacing harness service loops.

3.2.1.6 LandIS Access, Integration & Deintegration

SCIRD-182 The spacecraft shall allow mechanical integration and deintegration of the

LandIS without removal of spacecraft components, including spacecraft harnesses.

Rationale: Harnesses may be mated or demated as part of the integration and deintegration process.

SCIRD-219 The spacecraft shall provide unobstructed access to LandIS items that are to be installed, removed, or replaced after installation without disassembly.

SCIRD-220 The spacecraft shall mechanically integrate or deintegrate the LandIS vertically to the horizontal spacecraft interface.

Rationale: The LandIS should be lowered to the spacecraft interface. The +X Observatory Mechanical Body Frame axis is pointing upward from the ground.

3.2.1.7 LandIS Mounting

SCIRD-651 The spacecraft thruster placement and shielding shall be designed to minimize the radiative effects from a hot thruster on the LandIS volume.

SCIRD-187 The spacecraft shall accommodate the LandIS deck mounted with kinematic mounts.

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Rationale: LandIS deck is provided by the LandIS vendor. The baseline assumption is that the flight mounting flexures will be provided by the Instrument Vendor.

SCIRD-198 The spacecraft shall accommodate a LandIS mounting interface in the Y, Z plane, in the +X axis direction on the Spacecraft bus.

Note: The axes are defined in the SC mechanical body Reference frame.

Rationale: The LandIS will be mounted on the anti-ram face of Observatory orbital reference frame to reduce orbital debris impacts on the sensor.

3.2.1.8 LandIS Dynamics

SCIRD-200 Spacecraft hardware shall be designed to have positive margins of safety in all supporting structure and interfaces, including the LandIS interface, using the accelerations defined in the Mass Acceleration Curve (MAC) for components as shown in Figure 3.2.1.8-1, until the loads and accelerations are updated in the preliminary coupled loads analysis.

Rationale: These loads represent the quasi-static loads due to the low-frequency launch environment. The hardware should be designed for the loads given below as well as for loads induced by the high frequency random environment. The breakpoints for the MAC are given in Table 3.2.1.8-1. Design loads shown below should be updated based on the results of coupled loads analysis. Linear interpolation may be used between break-points to determine limit loads.

Note: The MAC is expected to be superseded by a combined launch vehicle and Observatory coupled loads analysis (CLA), and the CLA will presumably be the driver for Observatory-level dynamics testing.

Figure 3.2.1.8-1 Mass Acceleration Curve

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Table 3.2.1.8-1 Component Design Loads (Break Points)

Hardware Mass (kg) Quasi-static Load (g)

1.0 68.0

5.0 50.0

11.0 38.0

22.0 30.2

32.0 25.9

43.0 23.2

53.0 21.2

64.0 19.6

74.0 18.4

85.0 17.4

95.0 16.5

106.0 15.8

116.0 15.2

127.0 14.6

137.0 14.1

150.0 13.6

160.0 13.2

170.0 12.8

180.0 12.5

190.0 12.2

200.0 11.9

220.0 11.4

250.0 10.7

300.0 10.0

400.0 10.0

700.0 10.0

3.2.1.9 Instrument Mounted Spacecraft Components (IMSC) Sensor Accommodation SCIRD-205 Spacecraft sensors, such as gyros and star trackers, and associated mounting hardware and brackets, shall be mounted on the LandIS deck.

Rationale: To meet the pointing requirements.

SCIRD-214 The total IMSC volume, including all mounting hardware and brackets, shall not exceed the dimensions shown in the Figure shown below.

Rationale: The IMSC dimensions are not to exceed values. The IMSC volume is assumed to be bookkept within the LandIS volume. This volume does not include any Ground Support Equipment (GSE) necessary for ground testing.

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

Figure 3.2.1.9-1 IMSC Footprint

SCIRD-208 The spacecraft shall have an IMSC total mass not to exceed 50 Kg.

Rationale: Mass of IMSC, which reside on the LandIS instrument deck, are provided by SC.

This allocation is part of the overall SC budget and not part of the LandIS budget, and includes power and data harnessing. This allocation includes all hardware, secondary structure, harness, Multi-Layer Insulation (MLI) and Micrometeoroid and Orbital Debris (MMOD) blanketing.

SCIRD-316 The spacecraft IMSC shall be thermally isolated from the LandIS deck with a total linear conductance of less than or equal to 0.1 W/C. Note: Isolation (radiative and conductive) of the IMSC to the LandIS deck will be the responsibility of the spacecraft vendor and the spacecraft is responsible for the MLI closeouts between the IMSC and the LandIS deck.

Rationale: Heat transfer from the IMSC should not impact instrument thermal conditions.

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

SCIRD-714 The spacecraft shall accommodate a LandIS deck interface temperature to the spacecraft IMSC of 0°C to +30°C.

SCIRD-720 The spacecraft star trackers shall have individual unobstructed glint free zenith-directed fields of view out of the Y-Z plane of 25 deg along track by 25 deg cross-track, where the boresight of each star tracker FOV is clocked 45 degrees in opposite directions relative to the -Z axis (zenith).

Rationale: To meet mission pointing requirements, each of the two star trackers will be offset 90 degrees relative to one another to allow for simultaneous sensing of independent FOVs.

3.2.2 Orbit Control, Pointing and Alignment

The values contained within this section are consistent with the Geometric Error Budget (GEB), LNEXT-SYS-DESC-0016.

The requirements in this section will be iterated with the Spacecraft and LandIS teams upon selection.

Several TBRs contained in this section will be updated following line-of-sight jitter and ACS analyses. The process for developing the Finite Element Models (FEMs) is as follows: The instrument vendor is responsible for creating an Instrument Finite Element Model (FEM) based on their instrument design, and the spacecraft vendor will create a spacecraft FEM, based on their design. These are integrated by the spacecraft vendor to create an Observatory FEM. The Observatory FEM is used by the spacecraft vendor to perform analysis for instrument line-of-sight jitter generated by the spacecraft hardware, and by the instrument vendor to perform analysis for instrument line-of-sight jitter generated by instrument hardware.

SCIRD-254 The Landsat Next pointing reference shall be with respect to the Landsat

Next Orbital Reference Frame as defined in Section 3.2.1.1.

3.2.2.1 Pointing Stability and Jitter

Following the practice used on Landsat 9, Landsat Next will require management of the VSWIR line of sight motion due to jitter rather than specify a power spectral density of disturbance sources at the interface to the LandIS VSWIR. This approach requires the availability of mature LandIS finite element models for jitter analysis.

SCIRD-268 All Finite Element Models used for spacecraft and observatory jitter analysis shall incorporate a damping ratio of 0.25% with a Model Uncertainty Factor (MUF) of 1 or a damping ratio of 0.5% with a MUF of 2.

Rationale: A consistent and appropriate damping ratio will allow for consistent interpretation of modeling results.

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

The resulting predicted VSWIR line-of-sight disturbances in the along- and across-track directions must conform to the following requirements.

SCIRD-271 The effect of all spacecraft-generated translational acceleration and rotational disturbance sources (i.e., jitter), for all three axes, evaluated at the VSWIR line-of-sight shall not exceed 1.8 microradians, 3-sigma in the along- and across-track directions, for disturbance frequencies above 20 Hz.

Rationale: Jitter-induced line-of-sight variations contribute to geolocation and band registration errors, to the extent they are unmeasured, and to image smear, whether measured or not.

Disturbances at frequencies above 20 Hz are a particular concern, as they are difficult to measure using typical inertial reference units and are within the frequency range of expected LandIS structural modes, where flexible body effects make the corresponding line-of-sight transfer difficult to model and correct for, even if measurements are available.

3.2.2.2 Alignment

SCIRD-290 The spacecraft contractor shall measure Boresight Alignment Knowledge between the Spacecraft Attitude Determination Frame and the LandIS Interface (mounting surface datums) to an accuracy of 2 milliradians or less, 3-sigma, per axis, measured pre-launch.

Note: This measurement will be performed pre-launch by the spacecraft contractor and refined on orbit as part of the commissioning tests. The alignment between the LandIS boresight and the spacecraft attitude determination frame will be measured on orbit to much higher accuracy but pre-launch knowledge is required to be sufficient to initiate the on-orbit process to ensure that the alignment control requirements have been met.

SCIRD-296 The spacecraft shall provide alignment control between the Spacecraft

Attitude Determination Frame and the LandIS Interface (mounting surface datums) to an accuracy of 5 milliradians or less, 3-sigma, per axis.

Rationale: The LandIS boresight must be sufficiently aligned with the Spacecraft Attitude Determination Frame to limit the difference between the spacecraft ground track and the LandIS swath center to be less than the orbit control error (3-5 km) and to limit yaw-steering errors.

3.2.2.3 Spacecraft State Knowledge

SCIRD-294 Spacecraft Attitude Knowledge shall apply at the IMSC mounting location on the LandIS deck.

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

3.2.2.4 Disturbance Torques

SCIRD-483 The spacecraft shall accommodate periodic disturbance (with polarity change) torque magnitudes, including torques resulting from linear forces reacting from the LandIS to the spacecraft within the acceptable range of Figure 3.2.2.4-1 for all frequencies.

Figure 3.2.2.4-1 Torque Limit

SCIRD-715 The spacecraft shall accommodate disturbance torque with no polarity change to below 4x10-3 N-m for durations of 10 seconds or less and 1x10-3 N-m for durations above 10 seconds for each axis.

SCIRD-492 The on-orbit non-operational (non-imaging period) disturbances produced by the uncaging or deployment of any LandIS mechanism or device shall result in a torque, at the point of attachment with the properties defined in Figure 3.2.2.4-2.

Rationale: This requirement bounds the torques imparted on the spacecraft so that the attitude control system is able to recover attitude control after any LandIS disturbance caused by an uncaging or deployment event. These events will occur only during non-imaging periods.

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

Figure 3.2.2.4-2 Deployment Torque Limits

SCIRD-717 The spacecraft shall accommodate uncompensated momentum contributions from the LandIS of no more than +/- 0.5 N-m-sec per axis.

Rationale: LandIS uncompensated momentum will be accommodated by the Spacecraft.

3.2.3 Thermal Interface Requirements

3.2.3.1 Operating and Survival Heaters and Temperatures

The operating and survival temperatures, as well as the thermal isolation requirements specified in this section, are met at the mechanical interface between the spacecraft and the LandIS.

SCIRD-619 The spacecraft shall ensure that both the primary and redundant LandIS survival heater circuits are normally enabled on-orbit when the LandIS is off.

Survival heaters may be turned off in the event of an Observatory emergency where the survival of the spacecraft is in jeopardy.

SCIRD-723 The spacecraft interface survival temperature to the LandIS deck shall be -

20°C to +50°C.

SCIRD-724 The spacecraft interface operational temperature to the LandIS deck shall be

-10°C to +40°C.

SCIRD-725 The spacecraft interface operational diurnal temperature variation to the

LandIS deck shall be less than 10°C

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

3.2.3.2 Mechanical Mounting Interface Temperature Monitoring

SCIRD-312 The spacecraft shall monitor and report the temperatures of the spacecraft at the LandIS mechanical mounting interfaces in the spacecraft telemetry.

3.2.3.3 Thermal Isolation

SCIRD-721 The spacecraft shall accommodate a total linear conductance of less than or equal to 0.1 W/C from the LandIS deck to the spacecraft. Note: Isolation (radiative and conductive) of the LandIS deck to the spacecraft will be the responsibility of the LandIS vendor and the spacecraft vendor is responsible for the MLI closeouts between the IMSC and the LandIS deck.

3.2.4 Electrical Interface Requirements

3.2.4.1 Electrical Interfaces

SCIRD-320 The spacecraft shall provide the following electrical interfaces and services to the LandIS, by design.

-RS-422 command and telemetry data bus -Gigabit Ethernet (GbE) high-speed data bus -Operational Power Service -Survival Heater Power Service -Grounding Interface -1 Pulse Per Second (PPS) Interface -Discrete commands and telemetry, as needed -Discrete Bilevel & Analog telemetry, as needed

Rationale: Interfaces represent known LandIS interfaces and those that may be required (designated as “as needed”), pending finalization of the LandIS design. The Spacecraft will provide the instrument interface connector panel.

SCIRD-322 The spacecraft shall provide fault-protected, commandable power services, capable of selecting, connecting, and disconnecting non-critical LandIS components to/from the Main Bus.

Rationale: Defines basic parameters for switched power services.

SCIRD-324 The spacecraft shall provide Main Bus fault protection for LandIS-switched loads.

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

SCIRD-325 The spacecraft shall provide LandIS power service over-current fault protection circuits which are resettable by command.

SCIRD-327 The spacecraft shall provide LandIS Fault protected power services that survive a direct short circuit for any conditions prior to that fault clearing at the fault protection device.

Rationale: Ensures that power services are sized, and fault protection is set to ensure that a fault will clear at the intended fault protection site and that power service is viable, if reset.

SCIRD-331 The spacecraft power service wiring shall be designed to route parallel power and return wires in the same twisted groups.

Rationale: Minimizes Electromagnetic Interference (EMI) and magnetic dipole effects due to load power variations on the Main Bus.

SCIRD-337 The spacecraft shall provide switched power services having controlled rise times of between 0.5 and 3 ms, where rise time is defined as the time for the specified voltage to increase from 10% of final value to 90% of final value.

Rationale: This requirement bounds rise time to ensure clean turn on for loads while minimizing Ldi/dt that can affect turn-on transients and power quality.

3.2.4.2 Power Allocation

3.2.4.2.1 Orbit Average Power Allocation

SCIRD-342 The spacecraft shall provide an orbital average power to LandIS of no less than 375 W during LandIS imaging configuration at End of Life (EOL).

Rationale: This does not include the power required for any IMSC.

SCIRD-748 The spacecraft shall provide an orbital average power to LandIS of no less than 275 W when LandIS is in a survival configuration at End of Life (EOL).

Rationale: This does not include any power required for any IMSC.

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

3.2.4.3 Main Bus Power

3.2.4.3.1 Main Bus Voltage Range

SCIRD-351 The spacecraft shall provide LandIS survival heaters a minimum survival voltage of 24 V measured at the LandIS Main Bus Survival Heater power service interface connector panel between the spacecraft and LandIS.

Rationale: Ensure sufficient thermal control at a reduced voltage when the observatory is in a compromised state on-orbit that may prevent optimal power collection. The minimum Main bus voltage accounts for a least a single cell failure.

SCIRD-355 The spacecraft shall provide LandIS a voltage range between +28 to +34

VDC as measured at the LandIS Main Bus Operational power service interface connector panel between the spacecraft and LandIS.

Rationale: This defines Main Bus operational voltage range.

SCIRD-363 The spacecraft shall under no circumstances deliver a Main bus voltage exceeding 40 VDC at any Main Bus power service interface.

Rationale: This prevents the loss or damage to flight hardware that may result from human error, Electrical Ground Support Equipment, or other anomalies including but not limited to events during test/ground processing, credible faults, power-switching, and fail-over scenarios.

3.2.4.3.2 Main Bus Power Quality

SCIRD-358 Under steady state conditions (no load switching and no-fault clearing), voltage ripple on the Main Bus (at the input side of load switches) shall be less than 1.5V peak-to-peak over the frequency range of 30 Hz to 150 kHz.

Rationale: This defines Main Bus power quality based on CS101 conducted susceptibility limits and frequency range consistent with CS101.

SCIRD-360 Under steady state conditions (no load switching and no-fault clearing), voltage ripple on the Main Bus (at the input side of load switches) shall be less than 1V peak-to-peak over the frequency range of 150 kHz to 200 MHz.

Rationale: This defines Main Bus power quality based on CS114 conducted susceptibility limits and frequency range consistent with CS114.

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

SCIRD-366 The spacecraft-provided Main Bus operational power fault protection circuits shall not trip for an inrush transient having the following characteristics:

-10X load peak current for ≤ 10 microseconds -300% load peak current for > 10 microsecond and ≤ 10 milliseconds -120% load peak current for > 10 milliseconds

Rationale: Ensures that fault protection will not trip inadvertently due to turn-on transient, potentially rending it difficult or impossible to turn on. Avoids damage to the power service due to the turn on transient. Minimizes impact on Main Bus power quality.

3.2.4.4 Power Distribution Interfaces

SCIRD-345 LandIS Main Bus operational power service interfaces and Main Bus

Survival Heater power service interface shall be sized such that peak load current derating requirements of Instruction for EEE Parts Selection, Screening, Qualification, and Derating, EEE-INST-002 are met over the operational voltage range, with any single conductor failed open.

Rationale: Satisfies derating requirements for bundled wire in a vacuum, allowing for a single failed conductor that will not be detectable after flight mate.

SCIRD-372 The spacecraft shall provide…