Attachment D - L10 Spacecraft Interface Requirements Document (SCIRD).pdf

Attachment No. D

Landsat 10 Spacecraft to LandIS Interface Requirements

Document (SCIRD)

RFP 80GSFC26R0015

National Aeronautics and Space Administration Goddard Space Flight Center

Landsat 10 SCIRD L10-SC-ICD-0002 Revision C ii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Landsat 10 Spacecraft Interface Requirements Document (SCIRD) Signature/Approval Page

Prepared by:

Electronic Signature in TDMS Approval Date in TDMS James Hritzo Date Landsat 10 Spacecraft Systems Engineer NASA/GSFC, Code 590

Approved by:

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

NASA/GSFC, Code 426

Ross Henry Date Landsat 10 Project Manager NASA/GSFC, Code 426

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

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

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

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

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

Questions or comments concerning this document should be addressed to:

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

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

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

Name/Org.

Due Date 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 Reserved

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

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

3.2.2.3 Spacecraft State Knowledge

3.2.2.4 Disturbance Torques

3.2.3 Thermal Interface Requirements

3.2.3.1 Operational / 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 Overview

3.2.4.2 Power Services

3.2.4.3 Power Allocation

3.2.4.4 Power Bus

3.2.4.5 Power Distribution Interfaces

3.2.4.6 Spacecraft Class R Grounding Interface

3.2.4.7 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 (TOD) Pulse

3.2.5.3 Discrete Commands

3.2.6 Contamination Control

APPENDIX A ABBREVIATIONS AND ACRONYMS

APPENDIX B DEFINITIONS

List of Figures Figure 2.4-1 Landsat 10 Mission Architecture Figure 3.2.1.1.2-1 Observatory Orbital Reference Frame Figure 3.2.2.4-1 Torque Limit Figure 3.2.2.4-2 Deployment Torque Limits Figure 3.2.5.1.2-1 LandIS Data Cube Details on Ethernet Lanes

List of Tables

Table 2.4.1-1 Landsat 10 Mission Parameters Table 3.2.1.7-1 Spacecraft Interface Build Tolerance Table 3.2.1.7-2 On Orbit Thermal Distortion Limits Table 3.2.4.2-1 Spacecraft to LandIS Power Service Table

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1.0 INTRODUCTION

1.1 PURPOSE

The Landsat 10 Spacecraft Interface Requirements Document (SCIRD) sets forth the initial interface requirements imposed between the spacecraft and the Landsat 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 10 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 10 Interface Requirements documents (SCIRD, LIRD) are controlled and maintained by the Landsat 10 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

- Section C2 - Connectors and Contacts, Section W1 - Wire and Cable

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

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

1.3.2 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 133.0-B-2 Space Packet Protocol Recommended Standard, Issue 2, Editorial Change 1

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

CCSDS 301.0-B-4 Time Code Formats Recommended Standard, Issue 4, Editorial Change 1

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

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

Document Number Title Revision

1.3.3 Project Reference Documents

Document Number Title L10-SC-ICD-0005 Landsat 10 Spacecraft (SC) to Landsat Instrument Suite

(LandIS) Interface Control Document (SC-LandIS ICD) L10-SYS-DESC-0016 Landsat 10 Geometric Error Budget (GEB) L10-LANDIS-ICD-0015 LandIS Simplified Mechanical Interface Control Drawing

(MICD)

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

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

2.0 MISSION OVERVIEW

2.1 MISSION STATEMENT

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

2.2 MISSION BACKGROUND

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

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

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

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

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

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2.3 MISSION OBJECTIVES

Landsat 10 has these major mission objectives:

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

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

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

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

2.4 MISSION IMPLEMENTATION

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

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

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

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

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Figure 2.4-1 Landsat 10 Mission Architecture

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

2.4.1 Space Segment

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

Table 2.4.1-1 Landsat 10 Mission Parameters

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

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

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

2.4.2 Ground Segment

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

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

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

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

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

2.4.3 Launch Segment

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

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

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

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3.0 INTERFACE REQUIREMENTS

3.1 GENERAL

The Landsat 10 Interface Requirements are maintained across two documents; namely the SCIRD and LandIS Interface Requirements Document (LIRD). They are controlled and maintained by the Landsat 10 Project office. The SCIRD is levied on the spacecraft contractor.

The LIRD is levied on the LandIS contractor. These IRDs are maintained at Level 3.5 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.

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 Observatory 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 LandIS. 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 10, the mechanical and orbital reference frames 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).

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.

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3.2.1.1.1.1 Observatory Mechanical Body Frame

The observatory body frame is a right-hand, orthogonal, body-fixed XYZ system with +Y / +Z plane at the SC to Payload Adapter Fitting (PAF) interface plane, and the +Z direction pointing to orbit nadir, the +X direction in the LV velocity vector (or up when sitting on the PAF / vertical config), and the +Y completing the right-handed, orthogonal coordinate system.

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 PAF interface plane, and the +Z direction pointing to orbit nadir, the +X direction in LV velocity vector (or up when sitting on the PAF / vertical config), and the +Y completing the right-handed, orthogonal coordinate system.

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 PAF / 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 the LandIS stowed, deployed and transient keep-out envelopes as defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

Rationale: The LandIS dynamic envelope includes all LandIS-provided hardware in the stowed configuration and incorporates allowances for MLI, fabrication tolerances, and launch-induced dynamic motion. Deployment sweep volumes represent the complete sweep path and the fully deployed positions of the deployable covers. The detailed Mechanical Interface Configuration Drawing will be provided upon spacecraft contract award.

3.2.1.3 LandIS Field of View

SCIRD-153 The spacecraft shall provide LandIS unobstructed and glint free optical and thermal Fields of View (FOVs) as defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

Rationale: The detailed Mechanical Interface Configuration Drawing will be made available after contract award.

3.2.1.4 Mass Properties

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

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SCIRD-160 The spacecraft shall accommodate a not-to-exceed (NTE) mass of 675 kg for LandIS that includes LandIS and all associated hardware, intra-LandIS harnessing, any spacecraft supplied harnessing supported on LandIS, IMSC, and project-controlled LandIS reserve mass.

Rationale: The NTE mass includes the IMSC mass is defined in SCIRD-208. Note that IMSC mass remains a spacecraft budgeted item.

SCIRD-169 The spacecraft shall accommodate LandIS having a stowed center of mass within the limits defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

Rationale: The IMSC in not included in the LandIS center of mass calculations.

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 access to the fasteners for LandIS installation as defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD). Note: The fasteners are installed from the LandIS side of the interface.

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.

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Rationale: The LandIS should be lowered to the horizontally oriented spacecraft interface (i.e., the +X Observatory Mechanical Body Frame axis is pointing upward from the ground during integration/deintegration).

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, and its plume such that it does not cause any operational temperature limit violations on the LandIS.

Rationale: Irradiance from thruster plume impingement and the thruster itself should not harm LandIS.

SCIRD-187 The spacecraft shall accommodate the LandIS mechanical mounting interfaces as defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

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.

SCIRD-760 The Spacecraft shall accommodate the LandIS that has a minimum frequency greater than 35 Hz when fixed at the LandIS to spacecraft interface.

Rationale: The spacecraft will need to provide sufficient stiffness at the LandIS interface to meet observatory modal frequency requirements. LandIS will have a first significant mode of no less than 35 Hz, when mounted on flexures to a rigid interface.

SCIRD-761 The Spacecraft shall limit the build tolerances at the LandIS interface to the values listed in Table 3.2.1.7-1.

Notes: (1) Reference L10-LANDIS-ICD-0015, LandIS Simplified MICD, for fitting locations; (2) Pin positions to be controlled by Government provided template; (3) Limits may be revisited after contract award.

Mutually agreed to values will be documented in the detail MICD.

Rationale: The LandIS mount is not fully kinematic at the spacecraft interface. Distortions due to thermal loads may cause high stresses or unacceptable distortions in the LandIS structure.

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Table 3.2.1.7-1 Spacecraft Interface Build Tolerance

SCIRD-771 The Spacecraft shall limit thermally induced distortions at the LandIS interface to the values listed in Table 3.2.1.7-2.

Notes: (1) Reference L10-LANDIS-ICD-0015, LandIS Simplified MICD, for fitting locations; (2) Pin positions to be controlled by Government provided template; (3) Limits may be revisited after contract award.

Mutually agreed to values will be documented in the detail MICD.

Table 3.2.1.7-2 On Orbit Thermal Distortion Limits

3.2.1.8 Reserved

3.2.1.9 Instrument Mounted Spacecraft Components (IMSC) Sensor Accommodation SCIRD-205 Any Instrument Mounted Spacecraft Components (IMSC) such as gyros, star trackers, and associated mounting hardware / brackets shall be mounted to the LandIS mounting interface as defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

Rationale: To meet the pointing requirements.

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SCIRD-214 The IMSC components, brackets, and mounting hardware shall fit in the volume defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

Note: Component baffles, test equipment, and harnesses are permitted to exceed the LandIS volume. Harnessing from the spacecraft to IMSC components may also exceed this volume, pending mutual agreement.

Closeout MLI may exceed the volume as mutually agreed to and documented in the MICD.

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

Rationale: The mass of the IMSC is allocated to the spacecraft mass budget rather than the LandIS budget. This mass allocation includes all associated hardware, secondary structure, power and data harnessing, Multi-Layer Insulation (MLI), and Micrometeoroid and Orbital Debris (MMOD) blanketing.

SCIRD-762 The IMSC assembly shall have a minimum modal frequency greater than 75

Hz when fixed at the LandIS mounting interface.

Rationale: The LandIS strongback modal frequency is sensitive to the stiffness of the IMSC assembly. 75 Hz was arrived at through iterative design and analysis cycles by the GSFC IMSC mass simulator team. LandIS contractor concurred with this value as it is sufficiently stiff to avoid coupling with the LandIS primary modes ~ 35 Hz.

SCIRD-316 The spacecraft IMSC shall be thermally isolated from the LandIS mounting interface with a total linear conductance of less than or equal to 0.1 W/ºC.

Note: The spacecraft contractor is responsible for providing radiative and conductive isolation between the IMSC and the LandIS strongback. The spacecraft is also responsible for all MLI closeouts at the interfaces between the IMSC and LandIS.

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

SCIRD-714 The mounting interface temperature on the LandIS side of the interface to the

IMSC shall be 0°C to +30°C during both LandIS operational and survival modes.

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SCIRD-720 Spacecraft star trackers, as part of the IMSC, shall have fields of view (FOVs) within the envelope defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD).

Note: The FOV envelope defined in L10-LANDIS-ICD-0015, LandIS Simplified Mechanical Interface Control Drawing (MICD) is glint-free and unobstructed by the LandIS.

Rationale: The FOV volume is notionally designed to provide sufficient clearance to allow two star trackers (clocked 45 degrees offset relative to the -Z axis (zenith)) to each maintain 25 degrees along and cross track FOVs on the Y-Z plane.

3.2.2 Orbit Control, Pointing and Alignment

The values contained within this section are consistent with the Geometric Error Budget (GEB), L10-SYS-DESC-0016.

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

The process for developing the Finite Element Models (FEMs) is as follows: The LandIS vendor is responsible for creating an LandIS Finite Element Model (FEM) based on their 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 LandIS line-of-sight jitter generated by the spacecraft hardware, and by the LandIS vendor to perform analysis for LandIS line-of-sight jitter generated by LandIS hardware.

SCIRD-254 The Landsat 10 pointing reference shall be with respect to the Landsat 10

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 10 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.

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

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

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. 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 strongback.

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.

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

Figure 3.2.2.4-1 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 spacecraft shall accommodate the LandIS generated torque disturbances with the properties specified in Figure 3.2.2.4-2 during on-orbit non-operational periods from the uncaging or deployment of any LandIS mechanism or device.

Rationale: This requirement bounds the torques imparted on the spacecraft so that the attitude control system can 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 LandIS less than or equal to +/- 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 Operational / 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 enabled by default when the LandIS is off.

Note: LandIS survival heater power may be turned off in the event of an observatory emergency where the survival of the observatory is in jeopardy.

SCIRD-723 The mounting interface temperature on the spacecraft bus side of the interface to the LandIS shall be -20°C to +50°C during LandIS survival mode.

SCIRD-724 The mounting interface temperature on the spacecraft bus side of the interface to the LandIS shall be -10°C to +40°C during LandIS operational mode.

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

SCIRD-725 The mounting interface temperature on the spacecraft bus side of the interface to the LandIS shall have a per orbit variation of less than 10°C during LandIS science operations.

3.2.3.2 Mechanical Mounting Interface Temperature Monitoring

SCIRD-312 The spacecraft shall monitor and report the temperatures at the LandIS mechanical mounting interfaces in 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/ºC from the LandIS mounting interface to the spacecraft.

Note: Isolation (radiative and conductive) of the LandIS mounting interface to the spacecraft will be the responsibility of the LandIS contractor and the spacecraft contractor is responsible for the MLI closeouts between the IMSC and the LandIS mounting interface.

3.2.4 Electrical Interface Requirements

3.2.4.1 Electrical Interfaces Overview

SCIRD-320 The spacecraft shall provide the following electrical interfaces and services to LandIS:

- Operational Power Services

- Survival Heater Power Services

- Grounding Interface

- High-Speed Mission Data Bus (3 Ethernet Lanes)

- Primary Telemetry & Command Data Bus [including TOD pulse and message] (RS-422)

- Redundant Telemetry & Command Data Bus [including TOD pulse and message] (RS-422)

- Discrete commands

- Temperature Sensors to be monitored by SC (Passive Analog Telemetry)

- Active and passive bi-level telemetry

Rationale: Represents known LandIS interfaces pending finalization of the LandIS design.

SCIRD-1054 Except for cross-strapped redundant power and signal services, the spacecraft shall route redundant electrical harness wires to LandIS in a physically separated harness bundle.

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

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.

3.2.4.2 Power Services

SCIRD-322 The spacecraft shall provide overcurrent fault-protected power services which are resettable by command to LandIS as defined in Table 3.2.4.1-1 that can connect and disconnect LandIS components to/from the power bus.

Rationale: Defines basic parameters for switched power services.

Table 3.2.4.2-1 Spacecraft to LandIS Power Service Table

Power Service #

Power Type

Prime/Red

LandIS Load Description

Peak Steady State

Current (A)

Resistance (Ohms)

Circuit Breaker Max Current (A)

1 Operational Primary 1 (Notes 1, 3)

Payload Electronics (including op heaters, motor encoder, motor, FPA)

11.21 N/A 15

2 Operational Primary 2 (Notes 1, 3)

Payload Electronics (including op heaters, motor encoder, motor, FPA)

11.21 N/A 15

3 Operational Redundant 1 (Notes 2, 4)

Payload Electronics (including op heaters, motor encoder, motor, FPA)

11.21 N/A 15

4 Operational Redundant 2 (Notes 2, 4)

Payload Electronics (including op heaters, motor encoder, motor, FPA)

11.21 N/A 15

5 Operational Primary Cryocooler Electronics + TMU power

12.11 N/A 15

6 Operational Redundant Cryocooler Electronics + TMU power

12.11 N/A 15

7 Survival Htr Primary CT TMU 1

(TIR) &

CT TMU 2

(TIR)

5.66 6.8 7.5

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

Peak Steady State

Circuit Breaker Max Current (A)

8 Survival Htr Primary MT TMU 1

(SWIR) &

MT TMU 2

(SWIR)

5.66 6.8 7.5

9 Survival Htr Primary TMU Radiator

1, TMU

Radiator 2, TMU

Radiator 3, TMU Radiator

4, TMU

Radiator 5

14.15 2.72 15

10 Survival Htr Primary Optical Bench 1.38 28 2.5 11 Survival Htr Primary Electronics

Radiator 1

2.83 13.6 5

12 Survival Htr Primary Electronics Radiator 2

2.83 13.6 5

13 Survival Htr Primary VNIR Radiator 1.16 33.1 2.5 14 Survival Htr Primary Space View

Cover Damper NADIR View Cover Damper

2.57 13.6 5

15 Survival Htr Redundant CT TMU 1

(TIR) &

CT TMU 2

(TIR)

5.66 6.8 7.5

16 Survival Htr Redundant MT TMU 1

(SWIR) &

MT TMU 2

(SWIR)

5.66 6.8 7.5

17 Survival Htr Redundant TMU Radiator

1, TMU

Radiator 2, TMU

Radiator 3, TMU Radiator

4, TMU

Radiator 5

14.15 2.72 15

18 Survival Htr Redundant Strongback 1.38 28 2.5 19 Survival Htr Redundant Electronics

Radiator 1

2.83 13.6 5

20 Survival Htr Redundant Electronics Radiator 2

2.83 13.6 5

21 Survival Htr Redundant VNIR Radiator 1.16 33.1 2.5

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

Peak Steady State

Circuit Breaker Max Current (A)

22 Survival Htr Redundant Space View Cover Damper NADIR View Cover Damper

2.57 13.6 5

23 Spare 15 24 Spare 15 25 Spare 10 26 Spare 10 27 Spare 7.5 28 Spare 7.5 29 Spare 5 30 Spare 5 31 Spare 2.5 32 Spare 2.5

1) LandIS will tie these services together internally. Spacecraft design should account for this on SC-side of

2) LandIS will tie these services together internally. Spacecraft design should account for this on SC-side of interface.

3) One power service for these interfaces is an option.

4) One power service for these interfaces is an option.

SCIRD-327 The spacecraft shall provide LandIS fault protected power services that survive a direct short circuit, with no damage or degradation.

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 power 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.

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 Power Allocation

3.2.4.3.1 Orbit Average Power Allocation

SCIRD-342 The spacecraft shall provide no less than 500 W orbital average power

(OAP) to LandIS during LandIS imaging configuration at end of life (EOL).

Rationale: This does not include the power required for any IMSC and should be verified by power budget analysis.

SCIRD-748 The spacecraft shall provide no less than 400 W orbital average power to

LandIS when LandIS is in a survival configuration at End of Life (EOL).

Rationale: This does not include any power required for any IMSC and should be verified by power budget analysis.

SCIRD-763 The Spacecraft shall accommodate peak power consumption by LandIS per operational mode as defined in L10-SC-ICD-0005, Spacecraft to LandIS Interface Control Document.

3.2.4.4 Power Bus

3.2.4.4.1 Power Bus Voltage Range

SCIRD-355 The spacecraft shall provide LandIS a voltage range between +27 to +35

VDC as measured at the LandIS power service interface connectors between the spacecraft and LandIS.

Rationale: This defines power bus operational voltage range.

SCIRD-363 The spacecraft shall provide built-in over-voltage protection such that the power bus voltage does not exceed 40 VDC at any LandIS power service

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.

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

3.2.4.4.2 Power Bus Power Quality

SCIRD-358 The spacecraft shall ensure that, under steady-state conditions (no load switching and no fault clearing), the voltage ripple on LandIS power services measured at the input side of load switches is less than 1.5 V peak-to-peak over the frequency range of 30 Hz to 150 kHz.

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

SCIRD-360 The spacecraft shall ensure that, under steady-state conditions (no load switching and no fault clearing), the voltage ripple on LandIS power services measured at the input side of load switches is less than 1.0 V peak-to-peak over the frequency range of 150 kHz to 200 MHz.

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

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

-10X load peak current for less than or equal to 10 microseconds -300% load peak current for > 10 microsecond and less than or equal to 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 Bus power quality.

3.2.4.5 Power Distribution Interfaces

SCIRD-345 The spacecraft shall size LandIS operational power service interfaces and survival heater power service interfaces such that the peak load current derating requirements of EEE-INST-002 (Instruction for EEE Parts Selection, Screening, Qualification, and Derating) are met across 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…