Attachment F - L10 Observatory Environment Req Document (ObsERD)R.pdf

Attachment No. F

Landsat 10 Observatory Environment Requirements

Document (ObsERD)

RFP 80GSFC26R0015

National Aeronautics and Space Administration Goddard Space Flight Center

Landsat 10 ObsERD L10-OBSV-REQ-0031 Revision -ii

Landsat 10 Observatory Environmental Requirements Document (ObsERD)

Signature/Approval Page Prepared by:

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

Approved by:

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

Ross Henry Date Landsat 10 Project Manager

Evan Webb Date Landsat 10 Mission Systems Manager NASA/GSFC, Code 533 iii

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

Change History Log v

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

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

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

1.0 GENERAL INFORMATION

1.1 Purpose

1.2 Mission Overview

1.2.1 Mission Statement

1.2.2 Mission Background

1.2.3 Mission Objectives

1.2.4 Mission Implementation

1.2.4.1 Space Segment

1.2.4.2 Ground Segment

1.2.4.3 Launch Segment

1.3 Applicability and Limitations

1.4 The Landsat 10 Verification Approach

1.5 Assurance Requirements

1.6 Responsibility for Administration

1.7 Landsat 10 Configuration Control and Distribution

1.8 Applicable and Reference Documents

1.8.1 Applicable Documents

1.8.1.1 Project Applicable Documents

1.8.2 Reference Documents

1.9 Criteria for Unsatisfactory Performance

1.10 Testing of Spare Hardware

1.10.1 "One-Shot" Items

1.11 Test Facilities, Calibration

1.12 Test Condition Tolerances

1.13 Test Measurement Considerations

2.0 VERIFICATION PROGRAM

2.1 System Performance Verification

2.2 General Verification Requirements

2.2.1 Test Sequence and Level of Assembly

2.2.2 Verification Program Tailoring

2.2.3 Qualification of Hardware by Similarity

2.2.4 Test Factors/Durations

2.2.5 Structural Analysis/Design Factors of Safety

2.3 Function and Performance Test Requirements

2.3.1 Hardware Performance Testing

vi

2.3.2 Long Duration and Failure Free System Level Test of Flight Software (FSW) 19

2.4 Structural and Mechanical Verification Requirements

2.4.1 Structural Loads Qualification

2.4.1.1 Quasi-Static Limit Loads

2.4.1.2 Structural Analysis

2.4.1.3 Modal Survey

2.4.1.4 Design Strength Qualification

2.4.1.5 Structural Reliability (Residual Strength Verification)

2.4.2 Vibroacoustic Qualification

2.4.2.1 Acoustic Testing

2.4.2.2 Observatory Vibroacoustic Qualification

2.4.2.3 Spacecraft Subsystem Vibroacoustic Qualification

2.4.2.4 Component/Unit Vibroacoustic Qualification

2.4.2.5 Retest of Reworked Hardware

2.4.2.6 Fatigue Life Considerations

2.4.3 Sinusoidal Sweep Vibration Qualification

2.4.3.1 Observatory Sine Sweep Vibration Tests

2.4.3.2 Subsystem and Component Sine Sweep Vibration Tests

2.4.4 Mechanical Shock Qualification

2.4.4.1 Observatory Mechanical Shock Tests

2.4.4.2 Subsystem Mechanical Shock Tests

2.4.5 Mechanical Function Verification

2.4.5.1 Life Testing of Mechanisms

2.4.5.2 Demonstration of Mechanical Function

2.4.5.3 Torque/Force Margin

2.4.5.4 Acceptance Requirements

2.4.6 Pressure Profile Qualification

2.4.7 Mass Properties Verification

2.4.7.1 Mass Properties Test Demonstration

2.5 Electromagnetic Compatibility (EMC) Requirements

2.5.1 General EMC Requirements

2.5.1.1 Requirements Applicability

2.5.1.2 Safety and Controls

2.5.1.3 Test Facility and Equipment Requirements

2.5.1.4 Class R Bonding Verification

2.5.1.5 Power Grounding Verification

2.5.2 Spacecraft (Bus and Components) EMC Requirements

2.5.2.1 Conducted Emissions Requirements

2.5.2.2 Conducted Susceptibility Requirements

2.5.2.3 Radiated Emissions (RE) Requirements

2.5.2.4 Radiated Susceptibility Requirements

2.5.2.5 Spacecraft Power Quality Requirements

2.5.2.6 Spacecraft Self-Compatibility

2.5.3 Observatory EMC Requirements

2.5.3.1 Observatory Power Quality Requirements

2.5.3.2 Observatory Radiated Emissions

vii

2.5.3.3 Observatory Radiated Susceptibility Requirements

2.5.3.4 Self-Compatibility

2.6 Thermal Verification Requirements

2.6.1 General Requirements

2.6.2 Unit Test

2.6.3 Instrument Test

2.6.4 Observatory Test

2.6.5 Flight Spares

2.6.6 Flight Unit Retest

2.6.7 Thermal Launch Environment

2.7 Cryogenic Requirements

2.7.1 Cryogenic Systems Vacuum and Thermal Verification Requirements

2.7.2 Cryo-Thermal-Vacuum Qualification

2.7.2.1 Cryo-Thermal Special Considerations

2.7.2.2 Cryo-Thermal Level of Testing

2.7.2.3 Cryo-Thermal Test Parameters

2.7.2.4 Cryo-Thermal Test Setup

2.7.2.5 Cryo-Thermal Demonstration

2.7.2.6 Cryo-Thermal Special Tests

2.7.3 Leak Checking

2.7.4 Thermal Balance Qualification

APPENDIX A ABBREVIATIONS AND ACRONYMS

APPENDIX B DEFINITIONS

List of Figures Figure 1.2.4-1 Landsat 10 Mission Architecture Figure 2.4.1.1-1 Design Load Factor (applied to Obs CG) Figure 2.4.1.1-2 Reference Mass Acceleration Curve Figure 2.4.2.2-1 LV Random Vibration MPE Design Envelope Figure 2.4.2.4.1-1 Reference for Generalized Random Vibration Test Levels Components

22.7-kg (50-lb) or less Figure 2.4.2.4.1-2 Component Minimum Workmanship Random Vibration Test Levels

45.5-kg (100-lb) or less Figure 2.4.2.4.1-3 Determination of Qualification and Acceptance Random Verification

Test Levels Figure 2.4.3.1.2-1 Maximum Equivalent Sine Levels Figure 2.4.4.1-1 Candidate LV Separation Shock MPE Design Envelope at the

Observatory to LV Interface Figure 2.4.4.2.3-1 Shock Response Spectrum (SRS) for assessing Component Test

Requirements Figure 2.5.3.2.1-1 Operational Radiated Electric Field Emissions Limits Figure 2.6.1-1 Notional Thermal Vacuum Test Temperature Profile

List of Tables

Table 1.2.4.1-1 Landsat 10 Mission Parameters viii

Table 1.12-1 Test Condition Tolerances Table 2.2.1-1 Flight System Hardware Levels of Assembly Table 2.2.4-1 Test Factors/Durations/Margins Table 2.2.5-1 Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads

Table 2.4-1 Structural and Mechanical Verification Test Requirements Table 2.4.1.1-1 Observatory CG Load Factors Table 2.4.1.1-2 Reference Mass Acceleration Curve Breakpoints Table 2.4.2.1-1 Candidate LV Payload Acoustics Design Envelope Table 2.4.2.2-1 Candidate LV Random Vibration MPE Levels at Spacecraft Interface Table 2.4.2.4.1-1 Reference for Generalized Random Vibration Test Levels Components

22.7-kg (50-lb) or less Table 2.4.2.4.1-2 Component Minimum Workmanship Random Vibration Test Levels

45.4-kg (100-lb) or less Table 2.4.3.1.2-1 Maximum Equivalent Sine Levels of Observatory Table 2.4.3.2-1 Reference Component/Subsystem Acceptance Preliminary Sine

Environment (All Axes) Table 2.4.4.1-1 Candidate LV Separation Shock MPE Design Envelope at the

Observatory to LV Interface Table 2.4.5.3-1 Torque/Force Margins By Project Phase Table 2.5.3.2.1-1 Observatory Aggregate Radiated Emissions Electric Field Limits at the

Launch Site Table 2.5.3.3.1-1 Launch Site Radiated Susceptibility Electric Field Exposure Table 2.5.3.3.1-2 Operational Radiated Susceptibility Electric Field Environment Table 2.6.1-1 Environmental Verification Test Level Temperature Ranges Table 2.6.2-1 Thermal Cycling Options for Mechanical and Propulsion Hardware Table 2.7.1-1 Cryogenic Vacuum, Thermal, Electrical and Mechanical Requirements

1.0 GENERAL INFORMATION

1.1 PURPOSE

This document provides requirements and guidelines for the environmental verification program for the Landsat 10 observatory and describes methods for implementing those requirements. It contains a baseline for demonstrating by test or analysis the satisfactory performance of hardware in the expected mission environments, and that minimum workmanship standards have been met. It elaborates on those requirements, gives guideline test levels, provides guidance in the choice of test options, and describes acceptable test and analytical methods for implementing the requirements. There will be environmental verification plans to address requirements specified within this document at the spacecraft bus and observatory levels.

This document, acronymized to the ObsERD, is directly based on the General Environmental Verification Standard (GEVS) for GSFC Flight Programs and Projects, GSFC-STD-7000. GEVS has been tailored to create the Landsat 10 Project-specific environmental verification requirements within this document. For the purposes of environmental verification of Landsat 10 Project flight hardware, GEVS is not a formally applicable document in its entirety, but instead, portions are referenced herein for numerous purposes, including the clarification of testing rationale and for specific test guidance.

1.2 MISSION OVERVIEW

1.2.1 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.

1.2.2 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.

1.2.3 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.

1.2.4 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.

Figure 1.2.4-1 Landsat 10 Mission Architecture

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

1.2.4.1 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 1.2.4.1-1 Landsat 10 Mission Parameters

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.

1.2.4.2 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.

1.2.4.3 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 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.

Section 1.2 was generated using Landsat 10 Common Boilerplate (L10-MGMT-DESC-0005) Rev E

1.3 APPLICABILITY AND LIMITATIONS

These requirements apply to Landsat 10 hardware and associated software that is to be launched on a Landsat 10 Project-designated launch vehicle.

These requirements apply to all space flight hardware (see definition of hardware, Appendix B), including ground support interface hardware that is developed as part of the Landsat 10 Project.

These requirements are written in accordance with the current GSFC practice of using a single protoflight observatory to qualify for space flight. The protoflight verification program, therefore, is given as the nominal test program for Landsat 10.

1.4 THE LANDSAT 10 VERIFICATION APPROACH

The Landsat 10 observatory, at the appropriate levels of assembly, must be verified under conditions that simulate the launch operations, flight operations and flight environment as realistically as possible.

Since testing at the component (or unit) level, or lower level of assembly for large components, often becomes a primary part of the verification program, all components should be operating and monitored during all environmental tests if practicable.

Environmental verification of hardware is only a portion of the total verification effort for Landsat 10, which establishes that the observatory will function correctly. The environmental test program provides confidence that the design will perform when subjected to environments more severe than expected during the mission and provides environmental stress screening to uncover workmanship defects.

The total verification process also includes the development of models representing the hardware, tests to verify the adequacy of the models, analyses, alignments, calibrations, functional/performance tests to verify proper operation, and finally end-to-end tests and simulations to show that the total system will perform as specified.

1.5 ASSURANCE REQUIREMENTS

In addition to the verification program, the assurance effort includes parts and materials selection and control, reliability assessment, quality assurance, software assurance, design reviews, and system safety per the Landsat 10 Spacecraft Statement of Work (SC SOW), L10-SC-SOW-0006

1.6 RESPONSIBILITY FOR ADMINISTRATION

The responsibility and authority for decisions in waving the requirements rest with the GSFC Landsat 10 Project Manager.

The requirements thus derived and deviations from the requirements of this document are subject to review and approval by the GSFC Landsat 10 Project.

1.7 LANDSAT 10 CONFIGURATION CONTROL AND DISTRIBUTION

This document is controlled and maintained by the GSFC Landsat 10 Project.

1.8 APPLICABLE AND REFERENCE DOCUMENTS

ObsERD-25 The Contractor shall use the version of all applicable documents in effect at the time of procurement throughout development, except as identified in specific requirement(s).

1.8.1 Applicable Documents

Document Number Title Revision Applicable Section(s)

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

A w/CHANGE

NASA-STD-5001 Structural Design and Test Factors of Safety for Space Flight Hardware

B w/CHANGE

Section 4.2.4

NASA-STD-5019 Fracture Control Requirements for Spaceflight Hardware

A w/CHANGE

FCR 6 through FCR

NASA-STD-8719.9 Lifting Standard C All NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload

Safety Requirements A All

Annex to NASA-STD- 8719.24

Annex to NASA Expendable Launch Vehicle Payload Safety Requirements:

Requirements Table

B w/ Change 1

MSFC-STD-3029 Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

A Tables II and III

MIL-HDBK-17-1 Department of Defense Handbook:

Composite Materials Handbook - Volume 1.

F Section 8.1.4

Document Number Title Revision Applicable Section(s)

Polymer Matrix Composites Guidelines for Characterizations of Structural Materials

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.

1.8.1.1 Project Applicable Documents

Document Number Title L10-SC-ICD-0002 Landsat 10 Spacecraft Interface Requirements Document

(SCIRD)

L10-SC-SOW-0006 Landsat 10 Spacecraft Statement of Work (SC SOW) L10-SYS-DESC-0040 Landsat 10 Radiated Susceptibility (RS) Survive-Without

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

1.8.2 Reference Documents

Document Number Title Revision NASA-HDBK-7004 Force Limited Vibration Testing D NASA-HDBK-7005 Dynamic Environmental Criteria A GSFC-STD-7000 General Environmental Verification Standard

(GEVS) for GSFC Flight Programs and Projects B

MIL-STD-461 Requirements for the Control of Electromagnetic Interference Characteristics of Subsystems and Equipment

G

MIL-STD-464 Electromagnetic Environmental Effects Requirements for Systems

D

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

1.9 CRITERIA FOR UNSATISFACTORY PERFORMANCE

Deterioration or any change in performance of any test item that does or could in any manner prevent the item from meeting its functional, operational, or design requirements throughout its mission is reason to consider the test item as having failed.

If during a test sequence a test item is operated in excess of design life and wears out or becomes unsuitable for further testing from causes other than deficiencies, a spare may be substituted.

ObsERD-3492 If the substitution of a test item that has exceeded its design life affects the significance of test results, the test during which the item was replaced and any previously completed tests that are affected shall be repeated to the extent necessary to demonstrate satisfactory performance.

1.10 TESTING OF SPARE HARDWARE

Reserved

1.10.1 "One-Shot" Items Reserved

1.11 TEST FACILITIES, CALIBRATION

Reserved

1.12 TEST CONDITION TOLERANCES

ObsERD-270 The values and measurement uncertainties in Table 1.12-1 shall be used unless a rationale is provided for using other test condition tolerances, subject to the approval of the Landsat 10 Project Office.

Table 1.12-1 Test Condition Tolerances

Acoustics Overall Level: ≤ 1 dB l/3 Octave Band Tol: Frequency (Hz) Tolerance (dB) f ≤ 40

40 < F < 3150 f ≥ 3150

+3, -6 ±3

+3, -6 Duration: +10%, -0% Antenna Pattern Determination ± 2 dB Electromagnetic Compatibility Voltage Magnitude: ± 5% of the peak value Current Magnitude: ± 5% of the peak value RF Amplitudes: ± 2 dB Frequency: ± 2 % Distance: ± 5% of specified distance or

± 5 cm, whichever is greater Humidity ± 5% RH Loads Steady-State

(Acceleration):

± 5%

Sine Burst Amplitude: ± 5% Static: ± 5% Magnetic Properties Mapping Distance

Measurement:

± 1 cm

Displacement of assembly center of gravity (c.g.)

from rotation axis:

± 5 cm

Vertical displacement of single probe centerline from c.g. of assembly:

± 5 cm

Mapping turntable angular displacement:

± 3 degrees

Magnetic Field Strength: ± 1 nT

Repeatability of magnetic measurements (short term):

± 5% or ± 2 nT, whichever is greater

Demagnetizing and Magnetizing Field Level:

±5% of nominal

Mass Properties Weight: ± 0.2% Center of Gravity: ± 0.15 cm (± 0.06 in.)

Moments of Inertia: ± 1.5% Mechanical Shock Response Spectrum: Frequency (Hz) Tolerance (dB) Simulated Fn ≤ 3 kHz

Fn ≥ 3kHz ± 6 +9/-6

Shaker Fn ≤ 3 kHz ± 3 Overall Spectrum > 50% of SRS magnitude above nominal test level

Time History: ± 10% Pressure Greater than 1.3e04 Pa

(Greater than 100 mm Hg):

± 5%

1.3e04 to 1.3e02 Pa (100 mm Hg to 1 mm Hg):

± 10%

1.3e02 to 1.3e01 Pa (1 mm Hg to 1 micron):

± 25%

Less than 1.3e01 Pa (less than 1 micron):

± 80%

Temperature ± 2C Vibration Sinusoidal: Amplitude

Frequency ± 10% ±2%

Random: RMS Level Accel. Spectral Density Duration

± 10% ± 3 dB + 10 % / - 0 %

1.13 TEST MEASUREMENT CONSIDERATIONS

2.0 VERIFICATION PROGRAM

2.1 SYSTEM PERFORMANCE VERIFICATION

This section applies to the observatory, including its constituents. The provisions apply to all flight hardware, and its associated software.

The hardware SOWs provide the guidelines for an environmental verification program and the CDRLs detail the many associated deliverables.

2.2 GENERAL VERIFICATION REQUIREMENTS

Sections 2.3, Function and Performance Test Requirements, through 2.7, Cryogenic Requirements, cover the specific requirements for implementing the Landsat 10 environmental verification program for verifying the observatory and its constituents.

2.3 Function & Performance Test Requirements

2.4 Structural and Mechanical

2.5 EMC

2.6 Thermal

2.7 Cryogenics

In general, the environmental tests are not performed at the spacecraft bus level, but they can be performed at lower levels when beneficial for risk reduction or other purposes. Functional and performance tests will generally be done at the spacecraft bus level prior to LandIS integration.

The basic provisions are written assuming protoflight hardware. They are, in general, also applicable to prototype hardware. Acceptance requirements are also given for the flight acceptance of previously qualified hardware. This applies to follow-on hardware (multiple copies of the same item) developed for the program, or hardware (from another program) qualified by similarity.

2.2.1 Test Sequence and Level of Assembly

The verification activities herein are grouped by discipline; they are not all in a recommended sequence of performance. The recommended order of testing is electrical, mechanical and thermal at the end of the testing sequence. No specific environmental test sequence is required, but the test program will be arranged in a way to best disclose problems and failures associated with the characteristics of the hardware and the mission objectives.

Table 2.2.1-1, Flight System Hardware Levels of Assembly provides a hierarchy of levels of assembly for the flight hardware, with examples. These level designators are based on those used in the Space Systems Engineering Database developed by The Aerospace Corporation for the Air Force, and agreed to by NASA Headquarters, GSFC, and JPL. The ObsERD environmental test requirements generally start at the “unit” level and end at the “observatory” level.

Table 2.2.1-1 Flight System Hardware Levels of Assembly

2.2.2 Verification Program Tailoring

This document assumes that the observatory is of modular design and can be tested at the unit/component, subsystem/LandIS, and observatory/system (integrated spacecraft and LandIS) levels of assembly.

The Contractor will develop a verification program as stated in the Landsat 10 Spacecraft Statement of Work (SC SOW), L10-SC-SOW-0006 that satisfies the intent of the required verification program while taking into consideration the specific characteristics of the mission and the hardware. For example:

An observatory subsystem, or instrument, may be a functional subdivision of the observatory, but it may be distributed throughout the observatory rather than being a grouped physical entity.

In this case, the environmental tests, and associated functional tests, will be performed, at a minimum, at the component, subsystem, spacecraft, and observatory levels. Performance tests and calibrations may still be performed on the functional subsystem or instrument.

The physical size of the system may necessitate testing at other levels of assembly. Facility limitations may not allow certain environmental tests to be performed at the observatory level. In this case, testing should be performed at the highest practicable level. Also, for very large subsystems/instruments, tests at additional levels of assembly may be added in order to adequately verify the hardware design, workmanship and/or performance.

In some cases, because of the hardware configuration, it may be reasonable to test more than one component at a time. The components may be stacked in their flight configuration and may therefore be tested as a "section." Part of the decision process must consider the physical size and mass of the hardware. The test configuration must allow for adequate dynamic or thermal stress inputs to the hardware to uncover design errors and workmanship flaws.

Some test requirements stated as subsystem/instrument requirements may be satisfied at a higher level of assembly if approved by the GSFC Landsat 10 Project. For example, externally induced mechanical shock test requirements may be satisfied at the system level by firing the environment-producing pyro. A simulation of this environment is difficult, especially for large subsystems or instruments.

Aspects of the design and/or mission may negate certain test conditions to be imposed. For example, if the on-orbit temperature variations are small, less than 5°C, the project may consider alternate testing to the thermal-vacuum cycling at the system, or instrument, level of assembly and increasing the hot and cold dwell times.

The same process must be applied when developing the test plan for an instrument. While testing is required at the instrument component and fully integrated instrument levels of assembly, additional test levels may be necessary because of hardware complexity or physical size.

2.2.3 Qualification of Hardware by Similarity

There are cases in which hardware qualified for one flight program is to be built and used on another program. Hardware that has been previously qualified may be considered qualified for use on a new program by showing that the hardware is sufficiently similar to the original hardware and that the previous qualification program has adequately enveloped the new mission environments. The details for performing this comparison should be defined by the Contractor with approval from the Landsat 10 Project. As a minimum the following areas should be reviewed and documented:

ObsERD-507 Design and test requirements of original hardware that was previously qualified shall be shown to envelope the design and test requirements of subsequent hardware that is to be qualified by similarity to include a detailed review of the test configuration and of all waivers and deviations that may have occurred during testing of the original hardware.

ObsERD-508 Manufacturing information of hardware that is to be qualified by similarity shall be reviewed to determine if changes have been made that would invalidate the previous hardware qualification, covering parts, materials, packaging techniques as well as changes to the assembly process or procedures.

ObsERD-509 Test experience with the previous flight build shall be reviewed to verify that no modifications were made to the hardware during testing to successfully complete the test program, or else to verify all changes to the heritage hardware as a result of testing have been implemented on the current flight hardware.

If the Landsat 10 Project review shows that the hardware is of sufficiently similar design as the first build and that the previous test requirements envelope the Landsat 10 specified environmental requirements, then the hardware can be, with Landsat 10 Project approval, treated as qualified and need only to be subjected to acceptance level test requirements. The review of the hardware for similarity must be documented and included as part of the Contractor end item verification package.

2.2.4 Test Factors/Durations

Test factors for prototype, protoflight, and acceptance are given in Table 2.2.4-1. Test Factors/Durations/Margins.

Table 2.2.4-1 Test Factors/Durations/Margins

Test Prototype

Qualification Protoflight

Qualification Acceptance

Structural Loads1 Level

Duration Centrifuge/Static Load4 Sine Burst

1.25 x Limit Load 1 minute

5 cycles @ full level per axis

1.25 x Limit Load 30 seconds

5 cycles @ full level per axis

1.0 x Limit Load 30 seconds

5 cycles @ full level per axis

Acoustics Level2 Duration

Limit Level + 3dB 2 minutes

Limit Level + 3dB 1 minute

Limit Level 1 minute

Random Vibration Level2 Duration

Limit Level + 3dB 2 minutes/axis

Limit Level + 3dB 1 minute/axis

Limit Level 1 minute/axis

Sine Vibration3 Level Sweep Rate

1.25 x Limit Level 2 oct/min

1.25 x Limit Level 4 oct/min

Limit Level 4 oct/min

Mechanical Shock Actual Device Simulated

2 actuations

1.4 x Limit Level

2 x Each Axis

2 actuations

1.4 x Limit Level

1 x Each Axis

1 actuations Limit Level

1 x Each Axis Thermal-Vacuum Max./min.

Allowable Flight Temperature (AFT).

± 10°C

Max./min. AFT ± 10°C

Max./min. AFT ± 5°C(vacuum)

EMC & Magnetics As Specified for Mission.

See Section 2.5

Same Same

1 - If qualified by analysis only, positive margins are to be shown for factors of safety of 2.0 on yield and 2.6 on ultimate as directed in Section 2.4.1.1, Quasi-Static Limit Loads.. Beryllium and composite materials may not be qualified by analysis alone as covered in Section 2.4.1.5, Structural Reliability (Residual Strength Verification). Test levels for weldments, beryllium, bonded and composite structure, including metal matrix, are to be 1.25 x Limit Level for acceptance testing as covered in Section 2.4.1.5.

2 - As a minimum, the test level should be equal to or greater than the workmanship level.

3 - The sweep direction must be evaluated and chosen to minimize the risk of damage to the hardware. If a sine sweep is used to satisfy the loads or other requirements, rather than to simulate an oscillatory mission environment, a faster sweep rate may be considered, e.g., 6-8 oct/min to reduce the potential for over stress.

4 - Shorter durations may be used in static testing if necessary to protect the hardware from damage due to facility limitations. If a shorter duration is used, then the dwell time at load must be sufficient to demonstrate that the target loading condition has been achieved within the specified tolerances, all test measurements have been recorded, and the structure is stable under the applied loading condition.

2.2.5 Structural Analysis/Design Factors of Safety

ObsERD-514 Structural and mechanical verification testing shall be supported by structural analysis to provide confidence that the hardware will not experience failure or detrimental permanent deformation under test or launch conditions.

ObsERD-3714 The factors of safety that shall be applied to limit loads in order to calculate structural margins are shown in Table 2.2.5-1, Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads.

These factors of safety have been selected to be consistent with the test factors shown in Table 2.2.4-1, Test Factors/Durations/Margins. The yield factor of safety ensures that a prototype or protoflight test can be conducted with low risk of the hardware experiencing detrimental yielding. The ultimate factor of safety provides adequate separation between yield and ultimate failure modes and ensures that the hardware will not experience an ultimate failure under expected loading conditions.

ObsERD-515 In the case of thermally induced loads or stresses, the factors of safety shown in Table 2.2.5-1, Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads for the static loading condition shall be used for calculating strength margins.

ObsERD-3715 If the absolute value of the temperature differential between the stressed and un-stressed condition for the hardware using flight acceptance temperatures, (as defined in Table 2.6.1-1, is equal to or greater than 20° C, then the static factors of safety from Table 2.2.5-1, Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads shall be applied to the loads/stresses induced by acceptance temperatures without test margin otherwise protoflight temperatures should be used.

NOTE: Both hot and cold conditions should be evaluated. Thermally induced load should be combined with mechanical loading due to launch loads, gravity, or external loads due to enforced deflection if these mechanical loads occur at the same time.

Table 2.2.5-1 Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads

Type Static Sine Random/Acoustic Metallic Yield 1.25 1.25 1.6

Metallic Ultimate 1.4 1.4 1.8 Stability Ultimate 1.4 1.4 1.8 Beryllium Yield 1.4 1.4 1.8

Beryllium Ultimate 1.6 1.6 2.0 Composite Ultimate 1.5 1.5 1.9

Bonded Inserts/Joints Ultimate 1.5 1.5 1.9

ObsERD-605 Factors of safety for pressurized systems shall be compliant with NASA

Expendable Launch Vehicle Payload Safety Requirements, NASA-STD-

8719.24 and chapters 11 & 12 of Annex to NASA Expendable Launch Vehicle Payload Safety Requirements: Requirements Table, NASA-STD-

8719.24 (Annex).

ObsERD-606 Factors of safety for glass and structural glass bonds shall be as specified in

Structural Design and Test Factors of Safety for Space Flight Hardware, NASA-STD-5001, Section 4.2.4.

ObsERD-608 The random/acoustic safety factors shown in Table 2.2.5-1, Flight Hardware

Design/Analysis Factors of Safety Applied to Limit Loads shall be applied to statistically derived peak response based on root mean square (RMS) level and calculated as a 3-sigma value as a minimum.

Factors shown in Table 2.2.5-1, Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads assume that qualification/protoflight testing for random and acoustic is performed at acceptance level plus 3dB. If the difference between acceptance and qualification levels is less than 3dB, then these factors may be applied to qualification level minus 3dB instead of analyzing to acceptance level.

2.3 FUNCTION AND PERFORMANCE TEST REQUIREMENTS

The SC SOW defines functional, performance and confidence tests that are meant to verify and validate the observatory's operation before, during, and after environmental testing. These tests along with all other calibrations, measurements/ demonstrations, alignments (and alignment verifications), end-to-end tests, simulations, etc., that are part of the overall verification program are to be described in the System Performance Verification Plan. The following requirements address the minimum times for performance tests, and for failure-free testing/simulations that are needed for the Landsat 10 observatory.

2.3.1 Hardware Performance Testing

ObsERD-623 One thousand (1000) hours of operating/power-on time shall be accumulated on all flight electronic hardware, including any spares, prior to launch.

NOTE: For electronics consisting of a prime and redundant sides, i.e. A and B sides, 1000 hours should be accumulated on each side.

NOTE: Additional operating hours will be required for LandIS at the observatory level. See below for the minimum hours will have accumulated prior to delivery to Observatory integration.

ObsERD-624 At the conclusion of the performance verification program, the observatory shall have demonstrated failure-free performance testing for at least the last 350 hours of operation via any combination of Comprehensive Performance Test (CPT), Limited Performance Test (LPT) and Mission Simulation testing.

NOTE: This addresses the required level of testing at the observatory level.

NOTE: Hardware changes during or after the verification program invalidate previous demonstration.

NOTE: Additional operating hours will be required for LandIS at the observatory level. See below for the minimum hours LandIS will have accumulated prior to delivery to Observatory integration.

ObsERD-2109 At the conclusion of the performance verification program, all flight electronic hardware shall have demonstrated failure-free performance testing for at least the last 350 hours of operation.

NOTE: This requirement addresses the need for each electronics box to have a specific number of hour, but not necessarily at the observatory level.

NOTE: The demonstration hours may be accumulated at the subsystem level or system level of assembly. Hardware changes during or after the verification program invalidate previous demonstration.

NOTE: Additional operating hours will be required for LandIS at the

ObsERD-2110 At the conclusion of the performance verification program, flight electronic hardware shall have demonstrated failure-free performance testing for at least 200 hours of operation during the thermal-vacuum test exposure, with 100 hours of the trouble-free operation being logged at the hot-dwell temperatures and 100 hours being logged at the cold-dwell temperature.

NOTE: Hardware changes during or after the verification program invalidate previous demonstration.

NOTE: Additional operating hours will be required for LandIS at the

The project best practice is to accumulate 1000 hours of operating time on all flight hardware, and to demonstrate trouble-free performance at high-, low-, and nominal temperature. However, it is understood that under certain conditions this goal may not be met. For example, hardware change-out, rework, or repair just prior to launch may not provide sufficient time to fulfill these requirements. Also, the retest requirements following component failure during system-level thermal vacuum, or other tests, must be evaluated on a case-by-case basis, taking into account the criticality of the hardware element and the risk impact on achieving mission goals, as technical justification for any waiver approval.

The LandIS instrument will accumulate at least 300 hours of operation per side, prior to shipment for integration to the observatory.

The LandIS instrument will accumulate at least 50 hours of failure-free operation, prior to shipment to the observatory for integration and test.

The LandIS instrument will accumulate at least 100 hours of operation in vacuum per side or any redundant hardware, prior to shipment for integration to the observatory.

2.3.2 Long Duration and Failure Free System Level Test of Flight Software (FSW)

2.4 STRUCTURAL AND MECHANICAL VERIFICATION REQUIREMENTS

A series of tests and analyses should be conducted to demonstrate that the flight hardware is qualified for the expected mission environments and that the design of the hardware complies with the specified verification requirements such as factors of safety, interface compatibility, structural reliability, workmanship, and associated elements of system safety.

ObsERD-632 Table 2.4-1, Structural and Mechanical Verification Test Requirements specifies the structural and mechanical verification activities that shall be completed at the observatory level of assembly.

Testing at lower levels of assembly as defined in Table 2.4-1, Structural and Mechanical Verification Test Requirements is shown for reference. In order to fully demonstrate compliance to verification requirements, it may be necessary to test at the unit and subsystem levels of assembly.

When the tests and analyses are planned, consideration must be given to the expected environments of structural loads, vibroacoustics, sine vibration, mechanical shock, and pressure profiles induced during all phases of the mission; for example, during launch, insertion into final orbit, preparation for orbital operations, and entry, descent, and landing. Verification must also be accomplished to ensure that the transportation and handling environments are enveloped by the expected mission environments. Mass properties and proper mechanical functioning should also be verified.

Table 2.4-1 Structural and Mechanical Verification Test Requirements

Requirement Observatory Subsystem/

Instrument Unit (Component)

Including Instrument Units (Components)

Structural Loads Modal Survey *

ObsERD-655 T2 *

Design Qualification * ObsERD-660

A,T/A1 *

Structural Reliability Primary & Secondary Structure

ObsERD-670

(A,T)1 *

Vibroacoustics Acoustics T T2 T2 Random Vibration T2 T2 T Sine Vibration T3,T4 T3

ObsERD-691 T3

ObsERD-692 Mechanical Shock T T5 T5 Mechanical Function A,T A,T - Pressure Profile - A,T2 A Mass Properties A/T A,T2 *

* = May be performed at observatory or component level of assembly if appropriate.

A = Analysis required.

T = Test required.

A/T = Analysis and/or test.

A, T/A1 = Analysis and Test or analysis only if no-test factors of safety given in Section 2.2.4-1, Test Factors/Durations/Margins, are used.

(A, T)1 = Combination of fracture analysis and proof tests on selected elements, with special attention given to beryllium, composites, bonded joints and weldments.

T2 = Test must be performed unless assessment justifies deletion.

T3 = Test performed to simulate low frequency transient vibration and any sustained periodic mission environment, or to satisfy other requirement such as strength qualification.

T4 = Test must be performed, if practicable, to simulate transient and any sustained periodic vibration mission environment.

T5 = Test required for self-induced shocks, but may be performed at observatory level of assembly for externally induced shocks.

Of equal importance for qualifying the hardware for expected mission environments are the testing for workmanship and structural reliability, which are intended to provide a high probability of proper operation during the mission. In some cases, the expected mission environment is rather benign and produces test levels insufficient to expose workmanship defects. The verification test must envelope the expected mission levels, with appropriate margins added for qualification, and impose sufficient stress to detect workmanship faults.

Specific structural reliability requirements regarding fracture control, beryllium structure, composite structure, bonded structural joints, and glass structural elements are given in Section 2.4.1.5, Structural Reliability (Residual Strength Verification).

The program outlined in Table 2.4-1, Structural and Mechanical Verification Test Requirements, assumes that the observatory is sufficiently modularized to permit realistic environmental exposures at the subsystem level. When that is not possible, or at the project's discretion, compliance with the subsystem requirements must be accomplished at a higher or lower level of assembly.